Compression in a gas turbine engine

ABSTRACT

A gas turbine engine for an aircraft comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions. A core to fan tip temperature rise ratio is in the range from 2.845 to 3.8.

CROSS-REFERENCE TO RELATED APPLICATION(S)

This application is a continuation of U.S. application Ser. No.17/697,630 filed Mar. 17, 2022, which is a continuation of U.S.application Ser. No. 17/345,588 filed Jun. 11, 2021, which is acontinuation of U.S. application Ser. No. 16/558,417 filed Sep. 3, 2019,which is based on and claims priority under 35 U.S.C. 119 from GreatBritain Application No. 1908972.1 filed on Jun. 24, 2019. The entirecontents of the above applications are incorporated herein by reference.

The present disclosure relates to a gas turbine engine for an aircraft,and more specifically to a gas turbine engine arranged to have specifiedrelative airflow temperatures at different locations when operating atcruise conditions.

Gas turbine engines for aircraft propulsion have many design factorsthat affect the overall efficiency and power output or thrust. A generalaim for a gas turbine engine is to provide thrust with low specific fuelconsumption (SFC). In order to reduce SFC during cruise conditions boththe thermal and propulsive efficiencies of the engine may be increased.

To enable a higher thrust at a high efficiency, a larger diameter fanmay be used. When making a larger engine however, simply scaling upcomponents of a known engine type may not provide a correspondingscaling of power/thrust and/or efficiency, for example due todifferences in heat transfer throughout the larger engine.Reconsideration of engine parameters and operating conditions maytherefore be appropriate in order to provide a low SFC.

As used herein, a range “from value X to value Y” or “between value Xand value Y”, or the likes, denotes an inclusive range; including thebounding values of X and Y. All temperatures and pressure referred toherein are total temperature or total pressure unless otherwise stated.Where an average temperature is referred to this is taken to be a meanvalue. All temperatures are in Kelvin unless otherwise stated.

According to a first aspect there is provided a gas turbine engine foran aircraft comprising: an engine core comprising a turbine, acompressor, a core shaft connecting the turbine to the compressor, andan annular splitter at which the flow is divided between a core flowthat flows through the engine core, and a bypass flow that flows along abypass duct, wherein stagnation streamlines around the circumference ofthe engine, stagnating on a leading edge of the annular splitter, form astreamsurface forming a radially inner boundary of a streamtube thatcontains all of the bypass flow; and a fan located upstream of theengine core, the fan comprising a plurality of fan blades extending froma hub, each fan blade having a leading edge and a trailing edge, whereina fan tip radius of the fan is defined between a centreline of theengine and an outermost tip of each fan blade at its leading edge and ahub radius is defined between and the centreline of the engine and anouter surface of the hub at the radial position of the leading edge ofeach fan blade, each fan blade having a radially outer portion lyingwithin the streamtube that contains the bypass flow. A fan rotor entrytemperature is defined as an average temperature of airflow across theleading edge of each fan blade at cruise conditions and a fan rotor exittemperature is defined as an average temperature of airflow across aradially outer portion of each fan blade at the trailing edge at cruiseconditions. A fan hub to tip ratio of:

$\frac{{the}{fan}{hub}{radius}}{{the}{fan}{tip}{radius}}$

is in the range from 0.2 to 0.285; and a fan tip temperature rise of:

$\frac{{the}{fan}{tip}{rotor}{exit}{temperature}}{{the}{fan}{rotor}{entry}{temperature}}$

is in the range from 1.11 to 1.05.

According to a second aspect, there is provided a gas turbine engine foran aircraft comprising: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor,wherein the engine core has a core radius defined between the centrelineof the engine and a forwardmost tip of the engine core; and a fanlocated upstream of the engine core, the fan comprising a plurality offan blades extending from a hub, each fan blade having a leading edgeand a trailing edge, wherein a fan tip radius of the fan is definedbetween a centreline of the engine and an outermost tip of each fanblade at its leading edge and a hub radius is defined between and thecentreline of the engine and an outer surface of the hub at the radialposition of the leading edge of each fan blade. A fan rotor entrytemperature is defined as an average temperature of airflow across theleading edge of each fan blade at cruise conditions and a fan tip rotorexit temperature is defined as an average temperature of airflow acrossa radially outer portion of each fan blade at the trailing edge atcruise conditions, wherein the radially outer portion of each fan bladeis or comprises the portion of each fan blade at a distance from thecentreline of the engine greater than the core radius. A fan hub to tipratio of:

$\frac{{the}{fan}{hub}{radius}}{{the}{fan}{tip}{radius}}$

is in the range from 0.2 to 0.285; and a fan tip temperature rise of:

$\frac{{the}{fan}{tip}{rotor}{exit}{temperature}}{{the}{fan}{rotor}{entry}{temperature}}$

is in the range from 1.11 to 1.05.

To achieve a high propulsive efficiency without compromising transferefficiency (how efficiently energy is transferred from the core streamto the bypass stream), the inventors appreciated that there should be arelatively low fan (and more specifically, fan tip) temperature rise. Arelatively low fan temperature rise may indicate that the fan is capableof high efficiency in terms of useful work done by the fan, inparticular avoiding energy wastage as temperature rise across the fanrelative to ideal isentropic compression. To achieve this highpropulsive efficiency, reducing fuel burn, a high-flow fan is desired;the fan is therefore arranged to have a low hub to tip ratio to increaseor maximise fan flow area for a given diameter.

An efficient aerodynamic fan design may therefore be provided to allow arelatively low hub to tip ratio and a relatively low fan temperaturerise—an engine cycle is selected to facilitate the fan temperature riseremaining within the specified range at cruise conditions. An efficientaerodynamic fan design may comprise, for example, one or more of (i) arelatively wide chord with a relatively long sweep, (ii) relatively lowsuction surface curvature, and (iii) a relatively low friction surface.

The skilled person would appreciate that specific fuel consumption(SFC), weight and drag combine to give “fuel burn” of an installedengine. Reducing the fan tip temperature rise below the range specifiedabove may require the use of an excessively large fan in order toachieve a required thrust level, resulting in undesirable increasedweight and installation constraints and potentially negating any SFCfuel burn benefits in the overall fuel burn of the engine when installedon an aircraft.

Reducing the hub to tip ratio below the range specified above maydeleteriously reduce fan strength. The skilled person would appreciatethat the fan root and disc are designed to be strong enough to supportthe fan blade tips under all loads likely to be experienced inoperation.

As compared to prior art engine designs, the engine as described hereinmay allow for one or more of reduced fuel burn, reduced noise, andreduced specific fuel consumption (SFC). The low hub to tip ratio,coupled with the low fan tip temperature rise, has been found to providea fuel burn efficiency improvement in various embodiments.

The fan hub to tip ratio may be in the range from 0.200 to 0.285, andoptionally in the range from 0.24 to 0.27.

The fan tip temperature rise may be equal to 1.1, and optionally equalto 1.11.

The gas turbine engine may further comprise a nacelle surrounding thefan and the engine core and defining a bypass duct outside of the enginecore. The fan tip rotor exit temperature and the fan rotor entrytemperature may each provide a temperature of airflow across the fanblade portion in a bypass stream of air about to enter the bypass duct.The radially outer portion of each fan blade may be, comprise, or form amajor part of the portion of each fan blade extending across theentrance to the bypass duct.

A high propulsive efficiency may be achieved by having a low specificthrust engine with a low fan pressure ratio. For example, a fan pressureratio is defined as the ratio of the mean total pressure of the air flowat the exit of the fan to the mean total pressure of the air flow at theinlet of the fan, and wherein, at cruise conditions:

-   -   the fan pressure ratio may be in a range between 1.2 and 1.45,        and optionally    -   the fan pressure ratio may be in a range between 1.35 and 1.43,        and further optionally    -   the fan pressure ratio may be 1.39.

The turbine may be a first turbine, the compressor a first compressor,and the core shaft a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

The Overall Pressure Ratio (OPR) at cruise may be greater than 40 andlower than 80, and optionally in the range from 45 to 55.

According to a third aspect there is provided a method of operating agas turbine engine on an aircraft, the gas turbine engine being asdefined in either of the preceding two aspects, wherein the methodcomprises operating the gas turbine engine to provide propulsion undercruise conditions such that the fan hub to tip ratio is in the rangefrom 0.2 to 0.285, and the fan tip temperature rise is in the range from1.11 to 1.05.

According to a fourth aspect there is provided a gas turbine engine foran aircraft comprising: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor,wherein a compressor exit temperature is defined as an averagetemperature of airflow at the exit from the compressor, the engine corefurther comprising an annular splitter at which the flow is dividedbetween a core flow that flows through the engine core, and a bypassflow that flows along a bypass duct and a fan located upstream of theengine core, the fan comprising a plurality of fan blades extending froma hub, each fan blade having a leading edge and a trailing edge.Stagnation streamlines around the circumference of the engine,stagnating on a leading edge of the annular splitter, form astreamsurface forming a radially inner boundary of a streamtube thatcontains all of the bypass flow. Each fan has a radially outer portionlying within the streamtube that contains the bypass flow. A fan rotorentry temperature is defined as an average temperature of airflow acrossthe leading edge of each fan blade at cruise conditions and a fan tiprotor exit temperature is defined as an average temperature of airflowacross a radially outer portion of each fan blade at the trailing edgeat cruise conditions.

A fan tip temperature rise is defined as:

$\frac{{the}{fan}{tip}{rotor}{exit}{temperature}}{{the}{fan}{rotor}{entry}{temperature}}.$

A core temperature rise is defined as:

$\frac{{the}{compressor}{exit}{temperature}}{{the}{fan}{rotor}{entry}{temperature}}.$

A core to fan tip temperature rise ratio of:

$\frac{{the}{core}{temperature}{rise}}{{the}{fan}{tip}{temperature}{rise}}$

is in the range from 2.845 to 3.8.

According to a fifth aspect there is provided a gas turbine engine foran aircraft comprising: an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor,wherein a compressor exit temperature is defined as an averagetemperature of airflow at the exit from the compressor, the engine corehaving a core radius defined between the centreline of the engine and aforwardmost tip of the engine core; and a fan located upstream of theengine core, the fan comprising a plurality of fan blades extending froma hub, each fan blade having a leading edge and a trailing edge. Aradially outer portion of each fan blade is or comprises the portion ofeach fan blade at a distance from the centreline of the engine greaterthan the core radius. A fan rotor entry temperature is defined as anaverage temperature of airflow across the leading edge of each fan bladeat cruise conditions and a fan tip rotor exit temperature is defined asan average temperature of airflow across a radially outer portion ofeach fan blade at the trailing edge at cruise conditions. A fan tiptemperature rise is defined as:

$\frac{{the}{fan}{tip}{rotor}{exit}{temperature}}{{the}{fan}{rotor}{entry}{temperature}}.$

A core temperature rise is defined as:

$\frac{{the}{compressor}{exit}{temperature}}{{the}{fan}{rotor}{entry}{temperature}}.$

A core to fan tip temperature rise ratio of:

$\frac{{the}{core}{temperature}{rise}}{{the}{fan}{tip}{temperature}{rise}}$

is in the range from 2.845 to 3.8.

The skilled person would appreciate that a high propulsive efficiencymay be achieved by having a low specific thrust engine with a low fanpressure ratio. To do this without compromising transfer efficiencythere should be a low fan tip temperature rise, which reduces energywastage as temperature rise across the fan relative to ideal isentropiccompression. To facilitate achieving low fuel burn, the gas turbineengine may require high thermal efficiency—this may be achieved byefficient core compression which is achieved by a high core temperaturerise.

The core to fan tip temperature rise ratio is therefore relatively highby virtue of the relatively high core temperature rise and relativelylow fan tip temperature rise. The engine cycle may be devised based onthese parameters.

Reducing the fan tip temperature rise below the range specified abovemay require the use of an excessively large fan, potentially resultingin undesirable increased weight and installation constraints andnegating any fuel burn benefits.

Increasing the core temperature rise beyond the range specified abovemay overheat engine materials, potentially weakening or damaging theengine, and/or may require more cooling air so reducing or negating anyefficiency benefit.

As compared to known engine designs, the engine as described herein mayallow for one or more of reduced fuel burn, reduced noise, and/orreduced specific fuel consumption. The combination of a high level ofcore temperature rise and a low fan tip temperature rise may provide animprovement in fuel burn efficiency by combining increased thermalefficiency and propulsive efficiency.

The core to fan tip temperature rise ratio may be in the range from2.845 to 3.800, and optionally in the range from 2.9 to 3.2.

The fan tip temperature rise may be in the range from 1.05 to 1.11.

The core temperature rise may be in the range from 3.1 to 4.0, andoptionally in the range from 3.3 to 3.5.

The engine may further comprise a nacelle surrounding the fan and theengine core and defining a bypass duct outside of the engine core. Thefan tip rotor exit temperature and the fan rotor entry temperature mayeach provide a temperature of airflow across the fan blade portion in abypass stream of air about to enter the bypass duct. The radially outerportion of each fan blade may be, comprise, or form a major part of theportion of each fan blade extending across the entrance to the bypassduct.

The engine may comprise more than one compressor. In such embodiments,the compressor exit temperature may be measured or defined at the exitfrom the highest pressure compressor.

The Overall Pressure Ratio (OPR) at cruise may be greater than 40 andlower than 80, and optionally in the range from 45 to 55.

The turbine may be a first turbine, the compressor a first compressor,and the core shaft a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

According to a sixth aspect, there is provided a method of operating agas turbine engine on an aircraft, the gas turbine engine being asdefined in either of the preceding two aspects, wherein the methodcomprises operating the gas turbine engine to provide propulsion undercruise conditions such that the core to fan tip temperature rise ratiois in the range from 2.845 to 3.8.

According to a seventh aspect there is provided a gas turbine engine foran aircraft comprising an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor,and a fan located upstream of the engine core, the fan comprising aplurality of fan blades extending from a hub, each fan blade having aleading edge and a trailing edge. A compressor exit temperature isdefined as an average temperature of airflow at the exit from thecompressor at cruise conditions and a core entry temperature is definedas an average temperature of airflow entering the engine core at cruiseconditions, and a core compressor temperature rise is defined as:

$\frac{{the}{compressor}{exit}{temperature}}{{the}{core}{entry}{temperature}}.$

The engine core further comprises an annular splitter at which the flowis divided between a core flow that flows through the engine core, and abypass flow that flows along a bypass duct. Stagnation streamlinesaround the circumference of the engine, stagnating on a leading edge ofthe annular splitter, form a streamsurface forming a radially innerboundary of a streamtube that contains all of the bypass flow. Each fanblade has a radially outer portion lying within the streamtube thatcontains the bypass flow. A fan rotor entry temperature is defined as anaverage temperature of airflow across the leading edge of each fan bladeat cruise conditions and a fan tip rotor exit temperature is defined asan average temperature of airflow across the radially outer portion ofeach fan blade at the trailing edge at cruise conditions. A fan tiptemperature rise is defined as:

$\frac{{the}{fan}{tip}{rotor}{exit}{temperature}}{{the}{fan}{rotor}{entry}{temperature}}.$

A core compressor to fan tip temperature rise ratio of:

$\frac{{the}{core}{compressor}{temperature}{rise}}{{the}{fan}{temperature}{rise}}$

is in the range from 2.67 to 3.8.

According to an eighth aspect, there is provided a gas turbine enginefor an aircraft comprising an engine core having a core radius definedbetween the centreline of the engine and a forwardmost tip of the enginecore, wherein the engine core comprises a turbine, a compressor, and acore shaft connecting the turbine to the compressor, the engine furthercomprising a fan located upstream of the engine core, the fan comprisinga plurality of fan blades extending from a hub, each fan blade having aleading edge and a trailing edge, wherein a radially outer portion ofeach fan blade is or comprises the portion of each fan blade at adistance from the centreline of the engine greater than the core radius.A compressor exit temperature is defined as an average temperature ofairflow at the exit from the compressor at cruise conditions and a coreentry temperature is defined as an average temperature of airflowentering the engine core at cruise conditions, and a core compressortemperature rise is defined as:

$\frac{{the}{compressor}{exit}{temperature}}{{the}{core}{entry}{temperature}}.$

A fan rotor entry temperature is defined as an average temperature ofairflow across the leading edge of each fan blade at cruise conditionsand a fan tip rotor exit temperature is defined as an averagetemperature of airflow across the radially outer portion of each fanblade at the trailing edge at cruise conditions and a fan tiptemperature rise is defined as:

$\frac{{the}{fan}{tip}{rotor}{exit}{temperature}}{{the}{fan}{rotor}{entry}{temperature}}.$

A core compressor to fan tip temperature rise ratio of:

$\frac{{the}{core}{compressor}{temperature}{rise}}{{the}{fan}{tip}{temperature}{rise}}$

is in the range from 2.67 to 3.8.

As discussed above for preceding aspects, a high propulsive efficiencyand high transfer efficiency may be achieved by having a low specificthrust engine with a low fan pressure ratio and a low fan tiptemperature rise. To provide a low fuel burn with this high propulsiveefficiency, a high thermal efficiency of the engine is also desirable.High thermal efficiency may be provided by a high core compressortemperature rise with a high level of efficient core compression.

The core compressor temperature rise is measured across the corecompressor(s) only, and not across the fan blade (by contrast, thetemperature change across the fan root is included in the coretemperature rise discussed in preceding aspects).

The core compressor to fan tip temperature rise ratio is thereforerelatively high by virtue of the relatively high core compressortemperature rise and relatively low fan temperature rise. The enginecycle may be devised based on these parameters.

Reducing the fan tip temperature rise below the range specified abovemay require the use of an excessively large fan, potentially resultingin undesirable increased weight and installation constraints andreducing or negating any fuel burn benefits.

Increasing the core compressor temperature rise beyond the rangespecified above may overheat engine materials, potentially weakening ordamaging the engine, and/or may require more cooling air so reducing ornegating any efficiency benefit.

As compared to prior art engine designs, the engine as described hereinmay allow for one or more of reduced fuel burn, reduced noise, and/orreduced specific fuel consumption. The combination of the high corecompressor temperature rise and low fan tip temperature rise may providean improvement in fuel burn efficiency by combining increased thermalefficiency and propulsive efficiency.

The core compressor to fan tip temperature rise ratio may be in therange from 2.67 to 3.7, and optionally in the range from 2.80 to 2.95.

The fan tip temperature rise may be in the range from 1.05 to 1.11.

The core compressor temperature rise may be in the range from 2.9 to4.0, and optionally in the range from 3.1 to 3.3.

The engine may further comprise a nacelle surrounding the fan and theengine core and defining a bypass duct outside of the engine core. Thefan tip rotor exit temperature and the fan rotor entry temperature mayeach provide an airflow temperature across the fan blade portion in abypass stream of air about to enter the bypass duct. The radially outerportion of each fan blade may be, comprise, or form a major part of theportion of each fan blade extending across the entrance to the bypassduct.

The engine may comprise more than one compressor. In such embodiments,the compressor exit temperature may be measured or defined at the exitfrom the highest pressure compressor.

The engine core may comprise a core casing arranged to separate a coreairflow within the casing from a bypass airflow outside the casing. Thecore entry temperature may be:

(i) the temperature of the core airflow at the radial position of theforwardmost point of the core casing;

(ii) the temperature of the core airflow at the radial position of theleading edge of the forwardmost rotor or stator of the (lowest pressure)compressor; and/or

(iii) the temperature of the airflow across the trailing edge of aradially inner portion of each fan blade, the airflow across theradially inner portion of each fan blade being arranged to provide thecore airflow. Temperatures (i) to (iii) may be at least substantiallyequal.

The Overall Pressure Ratio (OPR) at cruise may be greater than 40 andlower than 80, and optionally in the range from 45 to 55.

The turbine may be a second turbine, the compressor a second compressor,and the core shaft a second core shaft. The engine core may furthercomprise a first turbine, a first compressor, and a first core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

A core to fan tip temperature rise ratio of:

$\frac{{the}{core}{temperature}{rise}}{{the}{fan}{tip}{temperature}{rise}},$

as defined for the fourth to sixth aspects, may be in the range from2.845 to 3.8, and the optional features described for those aspects mayalso apply to the seventh and eighth aspects.

According to a ninth aspect, there is provided a method of operating agas turbine engine on an aircraft, the gas turbine engine being asdefined in the seventh or eighth aspects, wherein the method comprisesoperating the gas turbine engine to provide propulsion under cruiseconditions such that the core compressor to fan tip temperature riseratio is in the range from 2.67 to 3.8.

The method may further comprise operating the gas turbine engine toprovide propulsion under cruise conditions such that a core to fan tiptemperature rise ratio as defined for the fourth to sixth aspects may bein the range from 2.845 to 3.8, and the optional features described forthose aspects may also apply to the ninth aspect.

According to a tenth aspect there is provided a gas turbine engine foran aircraft comprising an engine core comprising a turbine, acompressor, and a core shaft connecting the turbine to the compressor,and a fan located upstream of the engine core, the fan comprising aplurality of fan blades extending from a hub, each fan blade having aleading edge and a trailing edge. A compressor exit temperature isdefined as an average temperature of airflow at the exit from thecompressor at cruise conditions and a core entry temperature is definedas an average temperature of airflow entering the engine core at cruiseconditions. A core compressor temperature rise is defined as:

$\frac{{the}{compressor}{exit}{temperature}}{{the}{core}{entry}{temperature}}.$

A fan rotor entry temperature is defined as an average temperature ofairflow across the leading edge of each fan blade at cruise conditions,and a fan root temperature rise is defined as:

$\frac{{the}{core}{entry}{temperature}}{{the}{fan}{rotor}{entry}{temperature}}.$

A core compressor to fan root temperature rise ratio of:

$\frac{{the}{core}{compressor}{temperature}{rise}}{{the}{fan}{root}{temperature}{rise}}$

is in the range from 2.76 to 4.1.

As discussed above, a high propulsive efficiency may be achieved byhaving a low specific thrust engine with a low fan pressure ratio. Thesame fuel burn considerations as discussed with respect to the aspectsdescribed above may also apply.

The inventors appreciated that arranging the engine such that thetemperature rise across the core compressor(s) is greater than thatacross the fan root may facilitate obtaining a low fuel burn gas turbineengine whilst maintaining fan operability. Most of the core temperaturerise may therefore be across the core compressor(s) rather than acrossthe fan root.

The fan root temperature rise is measured across the inner portion ofthe fan blades, for the gas stream entering the core, and not across theouter portions of the fan blades, for the gas stream entering the bypassduct, as is done for the fan tip temperature rise described in precedingaspects.

In some embodiments, there could be no temperature change across the fanroot such that the fan root temperature rise is equal to one. Astemperature will not decrease across the fan root in normal operation,the lowest value for the denominator of the core compressor to fan roottemperature rise ratio is therefore one, making the value of the corecompressor to fan root temperature rise ratio equal to the corecompressor temperature rise. Obtaining a core compressor temperaturerise greater than 4.1 is not expected with any current aerospacematerials, as higher temperatures may weaken or damage the engine,and/or may require more cooling air so negating any efficiency benefit.

The core compressor to fan root temperature rise ratio is thereforerelatively high by virtue of the relatively high core compressortemperature rise and relatively small change in temperature across thefan root. The engine cycle may be devised based on these parameters. Thecurvature of the fan root may be selected to provide a low fan roottemperature rise.

As compared to known engine designs, the engine as described herein mayallow for one or more of reduced fuel burn, reduced noise, and/orreduced specific fuel consumption. The combination of the high corecompressor temperature rise and low fan root temperature rise mayprovide an improvement in fuel burn efficiency by combining increasedthermal efficiency and propulsive efficiency.

The core compressor to fan root temperature rise ratio may be in therange from 2.76 to 4.10, and optionally in the range from 2.8 to 3.2.

The fan root temperature rise may be in the range from 1.03 to 1.09.

The core compressor temperature rise may be in the range from 2.9 to4.0, and optionally in the range from 3.1 to 3.3.

The engine core may have a core radius defined between the centreline ofthe engine and a forwardmost tip of the engine core and the fan rotorentry temperature may be defined as the average temperature of airflowacross a radially inner portion of the leading edge of each fan blade atcruise conditions. The radially inner portion of each fan blade may be,form a major part of, or comprise the portion of each fan blade at adistance from the centreline of the engine less than the core radius.

The engine may comprise more than one compressor. In such embodiments,the compressor exit temperature may be measured at the exit from thehighest pressure compressor.

The engine core may comprise a core casing arranged to separate a coreairflow within the casing from a bypass airflow outside the casing. Thecore entry temperature may be one or more of:

(i) the temperature of the core airflow at the radial position of theforwardmost point of the core casing;

(ii) the temperature of the core airflow at the radial position of theleading edge of the forwardmost rotor or stator of the (lowest pressure)compressor; and/or

(iii) the temperature of the airflow across the trailing edge of aradially inner portion of each fan blade, the airflow across theradially inner portion of each fan blade being arranged to provide thecore airflow.

The engine core may comprise an annular splitter at which the flow isdivided between a core flow that flows through the engine core, and abypass flow that flows along a bypass duct. Stagnation streamlinesaround the circumference of the engine, stagnating on a leading edge ofthe annular splitter, may form a streamsurface forming a radially outerboundary of a streamtube that contains all of the core flow. Each fanblade may have a radially inner portion lying within the streamtube thatcontains the core flow. The core entry temperature may be defined as anaverage temperature of airflow across the trailing edge of the radiallyinner portion of each fan blade at cruise conditions.

The curvature of the root portion of each fan blade may be less than thecurvature across the tip portion of the blade, for example being between40% and 60% less, and optionally around 50% less. The root portion maybe the radially inner portion of the blade as described elsewhereherein, and the tip portion may be the radially outer portion of theblade as described elsewhere herein.

The Overall Pressure Ratio (OPR) at cruise may be greater than 40 andlower than 80, and optionally in the range from 45 to 55.

The turbine may be a second turbine, the compressor a second compressor,and the core shaft a second core shaft. The engine core may furthercomprise a first turbine, a first compressor, and a first core shaftconnecting the first turbine to the first compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

A core to fan tip temperature rise ratio of:

$\frac{{the}{core}{temperature}{rise}}{{the}{fan}{tip}{temperature}{rise}},$

as defined for the fourth to sixth aspects, may be in the range from2.845 to 3.8, and the optional features described for those aspects mayalso apply to the seventh and eighth aspects.

A core compressor to fan tip temperature rise ratio of:

$\frac{{the}{core}{compressor}{temperature}{rise}}{{the}{fan}{tip}{temperature}{rise}}$

as defined for the seventh to ninth aspects, may be in the range from2.845 to 3.8, and the optional features described for those aspects mayalso apply to the tenth aspect.

According to an eleventh aspect, there is provided a method of operatinga gas turbine engine on an aircraft, the gas turbine engine being asdefined in the tenth aspect, wherein the method comprises operating thegas turbine engine to provide propulsion under cruise conditions suchthat the core compressor to fan root temperature rise ratio is in therange from 2.76 to 4.1.

The method may further comprise operating the gas turbine engine toprovide propulsion under cruise conditions such that a core to fan tiptemperature rise ratio as defined for the fourth to sixth aspects may bein the range from 2.845 to 3.8, and the optional features described forthose aspects may also apply to the eleventh aspect.

The method may further comprise operating the gas turbine engine toprovide propulsion under cruise conditions such that a core compressorto fan tip temperature rise ratio as defined in the seventh to ninthaspects is in the range from 2.67 to 3.8, and the optional featuresdescribed for those aspects may also apply to the eleventh aspect.

According to a twelfth aspect there is provided a gas turbine engine foran aircraft comprising: an engine core comprising a first turbine, afirst compressor, and a first core shaft connecting the first turbine tothe first compressor; and a second turbine, a second compressor, and asecond core shaft connecting the second turbine to the secondcompressor, the second turbine being a higher pressure turbine than thefirst turbine and the second compressor being a higher pressurecompressor than the first compressor, the engine core further comprisingan annular splitter at which the flow is divided between a core flowthat flows through the engine core, and a bypass flow that flows along abypass duct, wherein stagnation streamlines around the circumference ofthe engine, stagnating on a leading edge of the annular splitter, form astreamsurface forming a radially inner boundary of a streamtube thatcontains all of the bypass flow; and a fan located upstream of theengine core, the fan comprising a plurality of fan blades extending froma hub, each fan blade having a leading edge and a trailing edge, eachfan blade having a radially outer portion lying within the streamtubethat contains the bypass flow. A first turbine entrance temperature isdefined as an average temperature of airflow at the entrance to thefirst turbine at cruise conditions and a first turbine exit temperatureis defined as an average temperature of airflow at the exit from thefirst turbine at cruise conditions, and a low pressure turbinetemperature change is defined as:

$\frac{{the}{first}{turbine}{exit}{temperature}}{{the}{first}{turbine}{entrance}{temperature}}.$

A fan rotor entry temperature is defined as an average temperature ofairflow across the leading edge of each fan blade at cruise conditionsand a fan tip rotor exit temperature is defined as an averagetemperature of airflow across the radially outer portion of each fanblade at the trailing edge at cruise conditions and a fan tiptemperature rise is defined as:

$\frac{{the}{fan}{tip}{rotor}{exit}{temperature}}{{the}{fan}{rotor}{entry}{temperature}}.$

A turbine to fan tip temperature change ratio of:

$\frac{{the}{low}{pressure}{turbine}{temperature}{change}}{{the}{fan}{tip}{temperature}{rise}}$

is in the range from 1.46 to 2.0.

According to a thirteenth aspect, there is provided a gas turbine enginefor an aircraft comprising: an engine core having a core radius definedbetween the centreline of the engine and a forwardmost tip of the enginecore, the engine core comprising a first turbine, a first compressor,and a first core shaft connecting the first turbine to the firstcompressor; and a second turbine, a second compressor, and a second coreshaft connecting the second turbine to the second compressor, the secondturbine being a higher pressure turbine than the first turbine and thesecond compressor being a higher pressure compressor than the firstcompressor, and a fan located upstream of the engine core, the fancomprising a plurality of fan blades extending from a hub, each fanblade having a leading edge and a trailing edge, each fan blade having aradially outer portion defined as the portion of each fan blade at adistance from the centreline of the engine greater than the core radius.A first turbine entrance temperature is defined as an averagetemperature of airflow at the entrance to the first turbine at cruiseconditions and a first turbine exit temperature is defined as an averagetemperature of airflow at the exit from the first turbine at cruiseconditions, and a low pressure turbine temperature change is defined as:

$\frac{{the}{first}{turbine}{exit}{temperature}}{{the}{first}{turbine}{entrance}{temperature}}.$

A fan rotor entry temperature is defined as an average temperature ofairflow across the leading edge of each fan blade at cruise conditions,and a fan tip rotor exit temperature is defined as an averagetemperature of airflow across the radially outer portion of each fanblade at the trailing edge at cruise conditions. A fan tip temperaturerise is defined as:

$\frac{{the}{fan}{tip}{rotor}{exit}{temperature}}{{the}{fan}{rotor}{entry}{temperature}},$

and a turbine to fan tip temperature change ratio of.

$\frac{{the}{low}{pressure}{turbine}{temperature}{change}}{{the}{fan}{tip}{temperature}{rise}}$

is in the range from 1.46 to 2.0.

To achieve a high propulsive efficiency without compromising transferefficiency (how efficiently energy is transferred from the core streamto the bypass stream), the inventors appreciated that there should be alow fan tip temperature rise as discussed above.

The inventors appreciated that designing the engine to have a relativelylarge temperature change across the lower pressure turbine(s) may allowmore work to be extracted more efficiently. Increasing, or maximising,the temperature change across the low pressure turbine may thereforeprovide various advantages. The skilled person would appreciate that thetemperature change across the turbine(s), from front to back, isgenerally a fall in temperature—a larger, or increased, temperaturechange is therefore a larger fall or drop in temperature (a largermagnitude of change), which may be thought of as a more negativetemperature change (as compared to the temperature rises discussedelsewhere).

The turbine to fan tip temperature change ratio is therefore relativelyhigh by virtue of the relatively high low pressure turbine temperaturechange and relatively low fan tip temperature rise. The engine cycle maybe devised based on these parameters.

As compared to known engine designs, the engine as described herein mayallow for one or more of reduced fuel burn, reduced noise, and/orreduced specific fuel consumption. Engines according to this aspect mayhave a high transfer efficiency due to extracting work efficiently fromthe core stream using the low pressure turbine, and applying that usingan efficient, low temperature rise, fan.

The turbine to fan tip temperature change ratio may be in the range from1.5 to 1.8

The fan tip temperature rise may be in the range from 1.05 to 1.1, andoptionally may be equal to 1.11.

The low pressure turbine temperature change may be in the range from 1.6to 1.85, and optionally in the range from 1.65 to 1.8.

The engine may comprise more than two turbines. In such embodiments, thehighest pressure turbine of the engine may be selected as the secondturbine and the lowest pressure turbine of the engine may selected asthe first turbine.

The Overall Pressure Ratio (OPR) at cruise may be greater than 40 andlower than 80, and optionally in the range from 45 to 55.

The first (low pressure) turbine may have four or more rotor stages.

According to a fourteenth aspect, there is provided a method ofoperating a gas turbine engine on an aircraft, the gas turbine enginebeing as defined in the twelfth or thirteenth aspects, wherein themethod comprises operating the gas turbine engine to provide propulsionunder cruise conditions such that the turbine to fan tip temperaturechange ratio is in the range from 1.46 to 2.0.

According to a fifteenth aspect there is provided a gas turbine enginefor an aircraft comprising: an engine core comprising a first turbine, afirst compressor, and a first core shaft connecting the first turbine tothe first compressor; and a second turbine, a second compressor, and asecond core shaft connecting the second turbine to the secondcompressor, the second turbine being a higher pressure turbine than thefirst turbine and the second compressor being a higher pressurecompressor than the first compressor, and a fan located upstream of theengine core, the fan comprising a plurality of fan blades extending froma hub. A second turbine entrance temperature is defined as an averagetemperature of airflow at the entrance to the second turbine at cruiseconditions, a first turbine entrance temperature is defined as anaverage temperature of airflow at the entrance to the first turbine atcruise conditions, a second turbine exit temperature is defined as anaverage temperature of airflow at the exit from the second turbine atcruise conditions, and a first turbine exit temperature is defined as anaverage temperature of airflow at the exit from the first turbine atcruise conditions. A low pressure turbine temperature change is definedas:

$\frac{{the}{first}{turbine}{entrance}{temperature}}{{the}{first}{turbine}{exit}{temperature}},$

anda high pressure turbine temperature change is defined as:

$\frac{{the}{second}{turbine}{entrance}{temperature}}{{the}{first}{turbine}{entrance}{temperature}}.$

A low to high pressure turbine temperature change ratio of:

$\frac{{the}{low}{pressure}{turbine}{temperature}{change}}{{the}{high}{pressure}{turbine}{temperature}{change}}$

is in the range from 1.09 to 1.25.

The inventors appreciated that, to reduce fuel burn in a (optionallygeared) gas turbine engine with two turbines, there is an optimal leveland split of temperature rise and work between the two turbines. In suchan engine with two turbines, the first turbine may be a lower pressureturbine and be arranged to drive a core shaft, and thereby to drive thefan (optionally via a gearbox); the second turbine may be a higherpressure turbine and may be connected to a different, second, coreshaft. The higher pressure turbine may be replaced with multipleturbines in some embodiments.

To reduce fuel burn, and optionally reduce or minimise core size, and/orincrease or maximise thermal efficiency across the high pressureturbine, the inventors appreciated that a relatively low temperaturechange across the higher pressure turbine is beneficial.

As the fan in a high bypass ratio turbofan generally generates most ofthe thrust, the inventors appreciated that efficiency of energy transferfrom the low pressure turbine to the fan should be improved, and thatincreasing or maximising the temperature drop/change across the lowpressure turbine(s) (LPT) may therefore be beneficial, noting that theLPT in a geared engine is generally more efficient than the highpressure turbine (HPT). The temperature change across the LPT istherefore larger relative to that across the HPT.

The low to high pressure turbine temperature change ratio is thereforerelatively high by virtue of the relatively large low pressure turbinetemperature change and relatively small high pressure turbinetemperature change. The engine cycle may be devised based on theseparameters.

As compared to known engine designs, the engine as described herein mayallow for one or more of reduced fuel burn, reduced noise, and/orreduced specific fuel consumption.

Reduced fuel burn may therefore be achieved due to a combination of arelatively high thermal efficiency from improved loading of the turbinesand a relatively high transfer efficiency from improved loading of theLPT and increasing of the temperature change across it.

The low to high pressure turbine temperature change ratio may be in therange from 1.10 to 1.25.

The low pressure turbine temperature change may be in the range from 1.6to 1.85, and optionally in the range from 1.65 to 1.8.

The high pressure turbine temperature change may be in the range fromrange from 1.40 to 1.55, and optionally in the range from 1.44 to 1.52.

The first turbine may be arranged to receive airflow from the exit ofthe second turbine, such that the first turbine entrance temperature maybe at least substantially equal to the second turbine exit temperature(e.g. barring the effect of any introduced cooling air or the likesbetween the two).

The engine may comprise more than two turbines. The highest pressureturbine of the engine may be selected as the second turbine and thelowest pressure turbine of the engine may be selected as the firstturbine in such embodiments.

The engine may comprise:

(i) a total of two turbines, and the first turbine entrance temperaturemay be at least substantially equal to the second turbine exittemperature in such embodiments; or

(ii) more than two turbines, and the high pressure turbine temperaturechange may provide a measure of the temperature change across allturbines except the lowest pressure turbine in such embodiments.

The Overall Pressure Ratio (OPR) at cruise may be greater than 40 andlower than 80, and optionally in the range from 45 to 55.

The first (low pressure) turbine may comprise at least four rotorstages.

A turbine to fan tip temperature change ratio of:

$\frac{{the}{low}{pressure}{turbine}{temperature}{change}}{{the}{fan}{tip}{temperature}{rise}}$

as defined in the twelfth, thirteenth and fourteenth aspects may be inthe range from 1.46 to 2.0. The optional features described with respectto those aspects may also apply to the fifteenth aspect.

According to a sixteenth aspect, there is provided a method of operatinga gas turbine engine on an aircraft, the gas turbine engine being asdefined in the fifteenth aspect, wherein the method comprises operatingthe gas turbine engine to provide propulsion under cruise conditionssuch that the low to high pressure turbine temperature change ratio isin the range from 1.09 to 1.30.

The method may further comprise operating the gas turbine engine toprovide propulsion under cruise conditions such that the turbine to fantip temperature change ratio as defined in the twelfth, thirteenth andfourteenth aspects may be in the range from 1.46 to 2.0. The optionalfeatures described with respect to those aspects may also apply to thesixteenth aspect.

In any of the aspects described above, one or more of the followingfeatures may apply:

A specific thrust of the engine at cruise conditions, defined as netengine thrust divided by mass flow rate through the engine, may be inthe range from 50 to 100 Nkg⁻¹s, and optionally below 90 Nkg⁻¹s.

A quasi-non-dimensional mass flow rate Q may be defined as:

${Q = {W\frac{\sqrt{T_{0}}}{P_{0} \cdot A_{fan}}}},$

where:W is mass flow rate through the fan in Kg/s;T₀ is average stagnation temperature of the air at the fan face inKelvin;P₀ is average stagnation pressure of the air at the fan face in Pa;A_(fan) is the area of the fan face in m²;Q may have a value in the range from 0.025 to 0.038 Kgs⁻¹N⁻¹K^(1/2) atcruise conditions, and optionally in the range from 0.031 to 0.036Kgs⁻¹N⁻¹K^(1/2). Q may take a value less than or equal to 0.035Kgs⁻¹N⁻¹K^(1/2) at cruise conditions.

A fan tip loading at cruise conditions may be defined as dH/U_(tip) ²,where dH is the enthalpy rise across the fan (23) and U_(tip) is the(translational) velocity of the fan tip (68) is in the range from 0.25to 0.4, and optionally from 0.28 to 0.34, and wherein further optionallythe fan tip loading takes a value in the range from 0.29 to 0.31 atcruise conditions.

Cruise conditions may mean the conditions at mid-cruise of an aircraftto which the engine is attached, and optionally may mean the conditionsexperienced by the aircraft and engine at the midpoint between top ofclimb and start of descent.

The forward speed of the gas turbine engine at cruise conditions may bein the range of from Mn 0.75 to Mn 0.85, and, optionally, the forwardspeed of the gas turbine engine at cruise conditions may be Mn 0.8.

The cruise conditions may correspond to atmospheric conditions definedby the International Standard Atmosphere at an altitude of 11582 m and aforward Mach Number of 0.8. Alternatively, cruise conditions maycorrespond to atmospheric conditions defined by the InternationalStandard Atmosphere at an altitude of 10668 m and a forward Mach Numberof 0.85. The cruise conditions may correspond to atmospheric conditionsat an altitude that is in the range of from 10500 m to 11600 m, andoptionally at an altitude of 11000 m.

The fan tip radius may be in the range from 110 cm to 150 cm; or in therange from 155 cm to 200 cm.

The gas turbine engine may further comprise a gearbox that receives aninput from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. Optionally, thegearbox may have a gear ratio in the range of from 3.2 to 5, furtheroptionally in the range from 3.2 to 3.8.

As noted elsewhere herein, the present disclosure may relate to a gasturbine engine. Such a gas turbine engine may comprise an engine corecomprising a turbine, a combustor, a compressor, and a core shaftconnecting the turbine to the compressor. Such a gas turbine engine maycomprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although notexclusively, beneficial for fans that are driven via a gearbox.Accordingly, the gas turbine engine may comprise a gearbox that receivesan input from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. The input tothe gearbox may be directly from the core shaft, or indirectly from thecore shaft, for example via a spur shaft and/or gear. The core shaft mayrigidly connect the turbine and the compressor, such that the turbineand compressor rotate at the same speed (with the fan rotating at alower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

The gearbox may be a reduction gearbox (in that the output to the fan isa lower rotational rate than the input from the core shaft). Any type ofgearbox may be used. For example, the gearbox may be a “planetary” or“star” gearbox, as described in more detail elsewhere herein. Thegearbox may have any desired reduction ratio (defined as the rotationalspeed of the input shaft divided by the rotational speed of the outputshaft), for example greater than 2.5, for example in the range of from 3to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1,3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratiomay be, for example, between any two of the values in the previoussentence. Purely by way of example, the gearbox may be a “star” gearboxhaving a ratio in the range of from 3.1 or 3.2 to 3.8. In somearrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.32. These ratios maycommonly be referred to as the hub-to-tip ratio. The radius at the huband the radius at the tip may both be measured at the leading edge (oraxially forwardmost) part of the blade. The hub-to-tip ratio refers, ofcourse, to the gas-washed portion of the fan blade, i.e. the portionradially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches),260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm(around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380 (around150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160inches) or 420 cm (around 165 inches). The fan diameter may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds), for examplein the range of from 240 cm to 280 cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cmto 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for examplein the range of from 1800 rpm to 2300 rpm, for example in the range offrom 1900 rpm to 2100 rpm. Purely by way of further non-limitativeexample, the rotational speed of the fan at cruise conditions for anengine having a fan diameter in the range of from 330 cm to 380 cm maybe in the range of from 1200 rpm to 2000 rpm, for example in the rangeof from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpmto 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.28, 0.29,0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4. Thefan tip loading may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.31, or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratiomay be in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of form 12 to 16, 13 to 15, or 13 to 14. The bypassduct may be substantially annular. The bypass duct may be radiallyoutside the core engine. The radially outer surface of the bypass ductmay be defined by a nacelle and/or a fan case.

The overall pressure ratio (OPR) of a gas turbine engine as describedand/or claimed herein may be defined as the ratio of the stagnationpressure upstream of the fan to the stagnation pressure at the exit ofthe highest pressure compressor (before entry into the combustor). Byway of non-limitative example, the overall pressure ratio of a gasturbine engine as described and/or claimed herein at cruise may begreater than (or on the order of) any of the following: 35, 40, 45, 50,55, 60, 65, 70, 75. The overall pressure ratio may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds), for example in the range offrom 50 to 70. By way of further example, the OPR at cruise may be inthe range from 45 to 65; optionally from 45 to 55; and furtheroptionally equal to or around 52.

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹s, 105 Nkg⁻¹s, 100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s or 80Nkg⁻¹s. The specific thrust may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds), for example in the range of from 80 Nkg⁻¹s to100 Nkg⁻¹s, or 85 Nkg⁻¹s to 95 Nkg⁻¹s. Such engines may be particularlyefficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). Purely by way of example, agas turbine as described and/or claimed herein may be capable ofproducing a maximum thrust in the range of from 330 kN to 420 kN, forexample 350 kN to 400 kN. The thrust referred to above may be themaximum net thrust at standard atmospheric conditions at sea level plus15 degrees C. (ambient pressure 101.3 kPa, temperature 30 degrees C.),with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 1800K to 1950K. The maximumTET may occur, for example, at a high thrust condition, for example at amaximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the industryas the “economic mission”) of an aircraft to which the gas turbineengine is designed to be attached. In this regard, mid-cruise is thepoint in an aircraft flight cycle at which 50% of the total fuel that isburned between top of climb and start of descent has been burned (whichmay be approximated by the midpoint—in terms of time and/ordistance—between top of climb and start of descent. Cruise conditionsthus define an operating point of the gas turbine engine that provides athrust that would ensure steady state operation (i.e. maintaining aconstant altitude and constant Mach Number) at mid-cruise of an aircraftto which it is designed to be attached, taking into account the numberof engines provided to that aircraft. For example where an engine isdesigned to be attached to an aircraft that has two engines of the sametype, at cruise conditions the engine provides half of the total thrustthat would be required for steady state operation of that aircraft atmid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide—in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach Number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be part of the cruise condition.For some aircraft, the cruise conditions may be outside these ranges,for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000 m to15000 m, for example in the range of from 10000 m to 12000 m, forexample in the range of from 10400 m to 11600 m (around 38000 ft), forexample in the range of from 10500 m to 11500 m, for example in therange of from 10600 m to 11400 m, for example in the range of from 10700m (around 35000 ft) to 11300 m, for example in the range of from 10800 mto 11200 m, for example in the range of from 10900 m to 11100 m, forexample on the order of 11000 m. The cruise conditions may correspond tostandard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 30 kN to 35 kN) at aforward Mach number of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582 m). Purely by way of further example, the cruiseconditions may correspond to an operating point of the engine thatprovides a known required thrust level (for example a value in the rangeof from 50 kN to 65 kN) at a forward Mach number of 0.85 and standardatmospheric conditions (according to the International StandardAtmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

According to an aspect, there is provided an aircraft comprising a gasturbine engine as described and/or claimed herein. The aircraftaccording to this aspect is the aircraft for which the gas turbineengine has been designed to be attached. Accordingly, the cruiseconditions according to this aspect correspond to the mid-cruise of theaircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gasturbine engine as described and/or claimed herein. The operation may beat the cruise conditions as defined elsewhere herein (for example interms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine as described and/or claimedherein. The operation according to this aspect may include (or may be)operation at the mid-cruise of the aircraft, as defined elsewhereherein.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4A is a close up sectional side view of an upstream portion of thegas turbine engine shown in FIG. 2 , with indications of where varioustemperatures are to be measured marked;

FIG. 4B is the close up sectional side view of FIG. 4A with regionswithin which the various temperatures may be measured marked;

FIG. 5A is a sectional side view of the gas turbine engine shown in FIG.1 , with indications of where various temperatures are to be measuredmarked;

FIG. 5B is the sectional side view of FIG. 5A with regions within whichthe various temperatures may be measured marked;

FIG. 6 is a schematic side view of a gas turbine engine;

FIG. 7 is the schematic side view of a gas turbine engine as shown inFIG. 1 , with turbine details highlighted;

FIG. 8 illustrates a method:

FIG. 9 is a close up sectional side view of an upstream portion of thegas turbine engine shown in FIG. 2 , with indications of flows and areasmarked; and

FIG. 10 is a perspective view of an aircraft with two gas turbineengines mounted thereon.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In the embodiment being described, the nacelle inner radius at the axialposition of the leading edge blade tips 68 a is arranged to be slightlylarger than the fan tip radius 102, such that the fan 23 can fit withinthe nacelle 21 without the blade tips 68 rubbing the nacelle 21. Moreparticularly, in the embodiment being described the engine 10 comprisesan engine fancasing 21 a adjacent the blade tips 68 a; the nacelle 21 ismounted on/around the engine fancasing 21 a such that the enginefancasing 21 a and the nacelle 21 form and surround an outer surface ofthe gas path though the engine 10. Fancasing inner radius at the axialposition of the leading edge blade tips 68 a is arranged to be slightlylarger than the fan tip radius 102, such that the fan 23 can fit withinthe engine fancasing 21 a without the blade tips 68 rubbing the fancasing 21 a. In some alternative embodiments, the blade tips 68 a may bearranged to rub the fancasing 21 a.

In the embodiments shown in the Figures, the engine fancasing 21 aextends only in the region of the fan 23. In alternative embodiments,the fancasing 21 a may extend rearwardly, for example to the axiallocation of a bypass duct outlet guide vane (OGV) 58.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2 . The low pressure turbine 19 (see FIG. 1 ) drives the shaft26, which is coupled to a sun wheel, or sun gear, 28 of the epicyclicgear arrangement 30. Radially outwardly of the sun gear 28 andintermeshing therewith is a plurality of planet gears 32 that arecoupled together by a planet carrier 34. The planet carrier 34constrains the planet gears 32 to precess around the sun gear 28 insynchronicity whilst enabling each planet gear 32 to rotate about itsown axis. The planet carrier 34 is coupled via linkages 36 to the fan 23in order to drive its rotation about the engine axis 9. Radiallyoutwardly of the planet gears 32 and intermeshing therewith is anannulus or ring gear 38 that is coupled, via linkages 40, to astationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

Each of the compressors provided in the gas turbine engine 10 (e.g. thelower pressure compressor 14 and the high pressure compressor 15)comprises any number of compression stages, for example multiplecompression stages. Each compression stage may comprise a row of rotorblades and a row of stator vanes that are axially offset from eachother. The fan 23 also provides compression of airflow, and so providesan additional compression stage separate from those of the low and highpressure compressors 14, 15. A compression stage number is defined asthe total number of compression stages provided by the fan 23 and theone or more compressors 14, 15 provided in the gas turbine engine 10. Inthe presently described embodiment, the compression stage number istherefore the sum of the compression stages provided in the low pressurecompressor 14, the high pressure compressor 15 and the fan 23.

In other embodiments, the compression stages provided in the compressors14, 15 of the gas turbine engine may not be axial compression stages. Insome embodiments, one or more radial compression stages may be providedin addition, or alternatively, to the axial compression stages providedin each compressor. For example, in one embodiment, the low pressurecompressor 14 and/or the high pressure compressor 15 may comprise one ormore axial compression stages (each formed by a row of rotor blades andstators) followed by a radial compression stage provided downstream ofthe axial compression stage or stages. In yet other embodiments, each ofthe compressors may comprise only radial compression stages. Thecompression stage number is defined as the total number of compressionstages, including both radial and axial compression stages (includingthe fan 23). In all of the embodiments described above, each radialcompression stage may comprise a centrifugal compressor.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3 . Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3 . There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2 . For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2 .

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area. Whilst the described example relates to aturbofan engine, the disclosure may apply, for example, to any type ofgas turbine engine, such as an open rotor (in which the fan stage is notsurrounded by a nacelle) or turboprop engine, for example. In somearrangements, the gas turbine engine 10 may not comprise a gearbox 30.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1 ), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

As noted above, downstream of the fan 23 the air splits into twoseparate flows: a first air flow A into the engine core 11 and a secondair flow B which passes through the bypass duct 22 to provide propulsivethrust. Referring to FIGS. 4A and 4B, the first and second airflows A, Bsplit at a generally annular splitter 70, for example at the leadingedge of the generally annular splitter 70 at a generally circularstagnation line. The splitter 70 is provided by a forwardmost portion ofthe core casing 11 a in the embodiments being described, and mayalternatively be referred to as a forwardmost tip 70 of the engine core11 in some embodiments.

A stagnation streamline 110 stagnates on the leading edge of thesplitter 70. The stagnation streamlines 110 around the circumference ofthe engine 10 form a streamsurface 110. All of the flow A radiallyinside this streamsurface 110 ultimately flows through the engine core11. The streamsurface 110 forms a radially outer boundary of astreamtube that contains all of the flow that ultimately flows throughthe engine core, which may be referred to as the core flow A. All of theflow B radially outside the streamsurface 110 ultimately flows throughthe bypass duct 22. The streamsurface 110 forms a radially innerboundary of a streamtube that contains all of the flow B that ultimatelyflows through the bypass duct 22, which may be referred to as the bypassflow B. The streamsurface 110, and correspondingly the streamtube, maybe defined at cruise conditions.

The flow at the fan exit that subsequently flows through the engine core11 may therefore be defined by a streamtube that extends from the fanexit to the engine core 11. Such a streamtube may be bounded by aradially outer surface that terminates at the splitter 70, i.e. by aradially outer surface that is formed by streamlines 110 that terminateat a stagnation point on the splitter 70. Such a streamtube may begenerally annular. For example a cross-section through such a streamtubemay be substantially annular at any given cross-section perpendicular tothe engine (rotational) axis 9.

Definitions of various temperatures, radii, and other parameters areprovided below for ease of reference.

Fan Tip Radius:

The radius 102 of the fan 23, also referred to as the fan tip radius102, or R_(fan tip), may be measured between the engine centreline 9 andthe tip 68 a of a fan blade 64 at its leading edge 64 a (in a radialdirection). The fan diameter (D) may simply be defined as twice theradius 102 of the fan 23.

In the embodiments being described, the fan tip radius 102 is in therange from 95 cm to 200 cm, or from 110 cm to 200 cm. In someembodiments, the fan tip radius is in the range from 95 cm to 150 cm orfrom 110 cm to 150 cm. In some alternative embodiments, the fan tipradius is in the range from 155 cm to 200 cm

In some embodiments, the fan diameter is in the range from 190 cm to 300cm, or 220 cm to 300 cm. In some alternative embodiments, the fandiameter is in the range from 310 cm to 400 cm.

The skilled person would appreciate that fan blades 64 may expand inoperation, and that the fan tip radius 102 under cruise conditions maybe slightly greater than the fan tip radius 102 measured when the fan 23is not in use. Fan tip radius 102 may be defined under cruiseconditions. However, the skilled person would appreciate that the changein fan tip radius 102 is generally small compared to the fan tip radiusand that the radius as measured when not in operation may be used.

Hub Radius:

The hub radius, R_(hub), is the (radial) distance 103 (in metres)between the centreline 9 of the engine 10 and the radially inner pointon the leading edge 64 a of the fan blade (i.e. of radially inner pointof the gas-washed surface of the fan blade)—this is equivalent to theradius of the hub 66 of the fan 23 at the point at which the leadingedge of each blade 64 extends therefrom.

Fan Area:

The fan face area, A_(fan), which may also be described as the flow areaof the fan, is defined as the annular area between fan blade tips 68 andthe hub 66 at the axial position of the fan blade leading edge tip 68 a.The fan face area is measured in a radial plane (i.e. a planeperpendicular to the engine axis 9 and containing radii of the engine atthe axial position of the plane). The skilled person will appreciatethat A_(fan) is at least substantially equivalent to the area of theannulus formed between the hub 66 of the fan 23 and the inner surface ofthe nacelle 21 immediately adjacent the leading edge blade tips (as theblade tip leading edges 64 a are arranged to lie very close to the innersurface of the nacelle 21—noting the above comments about the fancasing21 a) for the fan engine 10 being described, and is therefore equivalentto the fan face area minus the area taken by the hub 66.

As referred to herein, the flow area of the fan (A_(fan)) is defined as:

A _(fan)=π(R _(fan tip) ² −R _(hub) ²)

Where:

R_(fan tip) is the radius 102 (in metres) of the fan 23 at the leadingedge (i.e. at the tips 68 a of the leading edge of the fan blades 64);R_(hub) is the distance 103 (in metres) between the centreline of theengine and the radially inner point on the leading edge of the fan blade(i.e. of radially inner point of the gas-washed surface of the fanblade)—this is equivalent to the radius of the hub 66 of the fan 23 atthe point at which the leading edge of each blade 64 is connectedthereto, and may be referred to as the hub radius 103.

In various embodiments, the ratio of the radius of fan blade 64 at itshub 66 to the radius of the fan blade at its tip 68 may be less than0.33.

In the embodiment being described, the flow area is defined in a radialplane, and can therefore be calculated using the fan tip radius 102 andthe hub radius 103.

In the embodiments being described, the fan tip radius 102 is in therange from 95 cm to 200 cm, or from 110 cm to 200 cm. In someembodiments, the fan tip radius is in the range from 95 cm to 150 cm orfrom 110 cm to 150 cm. In some alternative embodiments, the fan tipradius is in the range from 155 cm to 200 cm

In some embodiments, the fan diameter (twice the fan radius 102) is inthe range from 190 cm to 300 cm, or 220 cm to 300 cm. In somealternative embodiments, the fan diameter is in the range from 310 cm to400 cm.

Q.

A quasi-non-dimensional mass flow rate, Q, may be defined as:

${Q = {W\frac{\sqrt{T_{0}}}{P_{0} \cdot A_{fan}}}},$

where:W is mass flow rate through the fan in Kg/s;T₀ is average stagnation temperature of the air at the fan face inKelvin;P₀ is average stagnation pressure of the air at the fan face in Pa;A_(fan) is the area of the fan face in m².

As referred to herein, the area of the fan face (A_(fan)) is defined as:

$A_{fan} = {\frac{\pi D^{2}}{4}\left( {1 - \left( \frac{h}{t} \right)^{2}} \right)}$

Where:

D is the diameter (in metres) of the fan at the leading edge (i.e. atthe tips of the leading edge of the fan blades);h is the distance (in metres) between the centreline of the engine andthe radially inner point on the leading edge of the fan blade (i.e. ofradially inner point of the gas-washed surface of the fan blade); andt is the distance (in metres) between the centreline of the engine andthe radially outer point on the leading edge of the fan blade (i.e.t=D/2)

A_(fan) may also be referred to as a fan flow area as it corresponds tothe gas-washed area of the fan (the blade-swept area outside of thehub). This may be equivalently represented as:

A _(fan)=π(R _(fan tip) −R _(hub) ²)

as described above.

At cruise conditions, the value of Q may be in the range of from: 0.0295to 0.0335; 0.03 to 0.033; 0.0305 to 0.0325; 0.031 to 0.032 or on theorder of 0.031 or 0.032 Kgs⁻¹N⁻¹K^(1/2). Thus, it will be appreciatedthat the value of Q may be in a range having a lower bound of 0.029,0.0295, 0.03, 0.0305, 0.031, 0.0315 or 0.032 and/or an upper bound of0.031, 0.0315, 0.032, 0.0325, 0.033, 0.0335, 0.034, 0.0345 or 0.035. Allvalues of Q referred to herein are provided in units of Kgs⁻¹N⁻¹K^(1/2).

Temperatures

All temperatures referred to herein are total temperatures; the sum ofstatic temperature plus velocity/kinetic energy effects. Totaltemperatures may also be referred to as stagnation temperatures. Alltemperature values are listed in Kelvin, unless otherwise stated and alltemperature ratios and rises are likewise calculated in Kelvin. Alltemperatures are defined at cruise conditions, as defined above. Inparticular, the ISA standards for cruise conditions may provide anindication of an expected ambient temperature. “Average” temperature isused to indicate a mean temperature.

In the embodiments being described, temperatures may be defined ormeasured at the mid-cruise aerodynamic design point, which is defined asMn 0.85, and an altitude of 10700 m (35,000 ft), and optionally moreparticularly of 10668 m, for engines 10 of the embodiments beingdescribed. The skilled person would appreciate that these cruiseconditions are provided by way of example only and may vary for engines10 of other embodiments. Under differing conditions, the absolutetemperature values may vary whilst the ratios remain within the rangesdescribed.

The following temperatures are referred to herein, and a more detaileddescription of each is provided below in Table 1. The numbering used forthe temperatures corresponds to that provided in SAE standard AS755F.

TABLE 1 Temperatures Approximate temperature in various embodiments(Kelvin) T120 the fan (tip) rotor 23 entry In the range 235 to 265;optionally in the range temperature 242 to 252 (approximately equalacross full May be, for example 244 or 250 blade length - thereforegenerally (altitude- and Mach number-dependent) equal to the fan rootrotor entry temperature T20) T125 the fan tip rotor 23 exit temperatureIn the range 260 to 285; optionally in the range (may vary along theblade length - 270 to 280 defined as an average across a May be, forexample 270 for a T120 of 244, or radially outer portion of the blade278 for a T120 of 250 unless otherwise specified) T30 the compressor 15exit temperature In the range 750 to 1050; optionally in the range (atthe exit from the highest pressure of 780 or 815 to 1000 compressor inembodiments with May be, for example, 834, 835 or 1000 multiplecompressors) T21 the core 11 entry temperature (may In the range 245 to270; optionally in the range be equivalent to a fan root 69 exit 260 to270 temperature) May be, for example, 260, 268 or 252 T42 a second(higher/highest pressure) In the range 960 to 1150; optionally in therange turbine 17 exit temperature T42 (also 1030 to 1100 or 1030 to1090. at least substantially equal to the first May be, for example,995, 1030 or 1100 (low pressure) turbine 19 entrance temperature T44 inthe embodiments described) T50 a first (low pressure) turbine 19 exit Inthe range 590 to 640; optionally in the range temperature 600 to 630 or605 to 615. May be, for example, 600, 612 or 630 T40 the second (highpressure) turbine 17 In the range 1400 to 1700; optionally in the entrytemperature range 1450 to 1650, or 1520 to 1570 May be, for example,1480, 1560 or 1650

For example, in one embodiment having a fan diameter in the range from330 to 380 cm, T120 may be 250 K and T125 may be 278 K, giving atemperature increase across the fan 23 of 28 Kelvin. The compressor exittemperature (T30) may be 834 K. The core entry temperature (T21) may be268 K. The second turbine exit temperature (T42) may be 1030 K and thefirst turbine exit temperature (T50) may be 612 K. The second turbineentry temperature (T40) may be 1560 K, giving a temperature decreaseacross the second turbine 17 of 530 K.

For example, in an alternative embodiment having a fan diameter in therange from 240 cm to 280 cm, T120 may be 245 K and T125 may be 270 K,giving a temperature increase across the fan 23 of 25 Kelvin. Thecompressor exit temperature (T30) may be 780 K. The core entrytemperature (T21) may be 260 K. The second turbine exit temperature(T42) may be 1000 K and the first turbine exit temperature (T50) may be630 K. The second turbine entry temperature (T40) may be 1480 K, givinga temperature decrease across the second turbine 17 of 480 K.

The skilled person would appreciate that one or more of the temperatureslisted in Table 1 may be measured or otherwise determined in variousways, for example by use of a temperature probe or rake, by modelling,or by indirect determination from a temperature measured (or otherwisedetermined) elsewhere in the engine 10. For example, T125 (fan tip exittemperature) may be measured by one or more probes mounted on a leadingedge of an outlet guide vane 59 in the bypass duct 22 (for example theclosest OGV 59 to the fan 23, if multiple bypass duct OGVs are present),or by a rake anywhere in the region labelled in FIG. 4B. Similarly, T21(fan root exit temperature) may be measured by one or more probesmounted on a leading edge of an outlet guide vane 24 in the core duct,or by a rake anywhere in the region labelled in FIG. 4B.

The skilled person would appreciate that one or more of the temperatureslisted in Table 1 may be difficult to measure practically, for examplethe relatively high temperature T40. Various temperatures may thereforebe inferred from temperature measurements elsewhere and a knowledge ofengine properties and temperature relationships.

The fan 23, which is located upstream of the engine core 11, comprises ahub 66 and a plurality of fan blades 64 extending from the hub 66. Eachfan blade 64 has a leading edge 64 a and a trailing edge 64 b.

The fan rotor entry temperature (T120) is defined as an averagetemperature of airflow across the leading edge 64 a of each fan blade 64at cruise conditions; in particular, the temperature may be definedacross a radially outer portion of the leading edge 64 a of each fanblade 64 at cruise conditions. The skilled person would appreciate thatthe temperature across the leading edge 64 a of each fan blade 64 may beat least substantially equal across the entire leading edge 64 a of thefan blade 64, and that an average across the outer portion (or acrossthe whole blade length) may be taken. More specifically, T120 may beused to refer to the temperature across a radially outer (fan tip)portion of the leading edge 64 a of each fan blade 64 at cruiseconditions and T20 may be used to refer to the temperature across aradially inner (fan root) portion of the leading edge 64 a of each fanblade 64 at cruise conditions. The term fan rotor entry temperature maytherefore be used generally for T120 or T20.

The fan rotor entry temperature T120 may be higher than the ambienttemperature, for example by around 30 K in some embodiments, due todynamic head/Mach number-related effects. The fan rotor entrytemperature (T120) may therefore be measured or calculated anywherewithin a relatively large region, as illustrated in FIG. 4B (anywherewithin the nacelle 21 and up to the fan blades 64, as the increase ofaround 30 K from ambient may occur at or near the forward-most edge ofthe nacelle 21), but may more specifically be measured at or adjacentthe fan's leading edge 64 a. The Fan Rotor Entry Temperature, T120, isaround 250K in the embodiment being described. The skilled person wouldappreciate that this may vary in other embodiments, based on factorssuch as altitude of cruise.

A radially outer portion of each fan blade 64 is selected in line withstandards for the definition of T120, and for convenience of comparisonwith other temperatures as described below. The skilled person wouldappreciate that, for the engine 10 shown at cruise, T10 (temperature atthe nacelle's forward-most point) is at least substantially equal toT120 (temperature at the leading edge 64 a of an outer region of the fanblade 64/near the fan tip), as the Mach number based temperatureincrease takes effect from that point, and that T10 and T120 may also besubstantially equal to T20 (temperature at the leading edge 64 a of aninner region of the fan blade 64/near the fan root).

The radially outer portion of each fan blade 64 may be defined as theportion of each fan blade 64 washed by the bypass airflow B, which flowsaround the outside of the engine core 11 after passing the fan 23 (asopposed to the core airflow, A, which passes through the core 11). Thisbypass airflow B flows through the bypass duct 22 in the embodimentsdescribed herein.

In the embodiment being described, the engine core 11 has a core radius105 defined between the centreline 9 of the engine 10 and a forwardmosttip of the engine core 11; the forwardmost tip may be referred to as thesplitter 70, as it divides the core airflow A from the bypass airflow B.The radially outer portion of each fan blade 64 is generally the portionof each fan blade 64 at a radial distance from the centreline 9 of theengine 10 greater than the core radius 105.

The skilled person would appreciate that, in reality the streamsurface110 may slope and/or curve relative to the engine axis 9, such that someof the gas stream passing the fan blade 64 at a radial distance from thecentreline 9 of the engine 10 slightly less than or equal to the coreradius 105 may still enter the bypass stream B in some embodiments. Inthe embodiments being described, the slope and/or curvature of thestreamsurface 110 relative to the engine axis 9 is relatively small,such that using the radial position of the splitter 70 provides an atleast substantially equivalent temperature to using the streamsurface110, within measurement errors. The division at a set radial positionmay therefore provide an equivalent value which may be easier todetermine than streamtube shape in some scenarios.

The gas turbine engine 10 of the embodiment being described comprises anacelle 21 surrounding the fan 23 and the engine core 11 and defining abypass duct 22 outside of the engine core 11. The bypass airflow B flowsthrough the bypass duct 22 after leaving the fan 23 in the embodimentsbeing described. The radially outer portion of each fan blade 64 istherefore the portion of each fan blade 64 extending across the entranceto the bypass duct 22 in the embodiments being described.

The fan tip rotor exit temperature (T125) is defined as an averagetemperature of airflow over the radially outer portion of each fan blade64 at the trailing edge 64 b of each fan blade 64 at cruise conditions.The radially outer portion is as defined for the leading edge 64 a. T125therefore corresponds to the temperature of the bypass stream B onleaving the fan 23.

The fan tip rotor exit temperature (T125) and the fan rotor entrytemperature (T120) both therefore refer to airflow temperatures acrossthe fan blade portion which is located in a bypass stream of air B aboutto enter the bypass duct 22 (the radially outer fan blade portion).

The Fan Rotor Entry Temperature, T20, may also refer to airflowtemperatures across the leading edge of the fan blade portion which islocated in a core stream of air A about to enter the engine core 11 (theradially inner fan blade portion/the portion of the gas stream radiallyinward of the streamsurface 110), as the temperature is equivalentacross the leading edge 64 a of the blade 64. This is generally not thesame for the exit temperature T125, T21, as the temperature generallyvaries with radius across the trailing edge 64 b of each blade.

The skilled person would appreciate that airflow temperature generallyincreases across the fan 23 when at cruise conditions, as work is doneon the air by the fan blades 64 and some of this work generallymanifests as heat. The exit temperature T125, T21 is therefore generallyhigher than the entry temperature T120, T20.

The ratio of the fan tip rotor exit temperature T125 to the fan tiprotor entry temperature T120, T125/T120 may therefore be referred to asthe fan tip temperature rise. The fan tip temperature rise may bedefined as the average temperature rise across the fan rotor portion inthe bypass stream (B in FIG. 1 ). The fan tip temperature rise isgreater than one, and more specifically is in the range from 1.11 to1.05, in the embodiments being described. The temperature rise may bedefined as the ratio of the mean total temperature of the flow at thefan exit that subsequently flows (as flow B) around the outside of theengine core 11 to the mean total temperature at the inlet to the fan 23.

The fan tip temperature rise is relatively low in the embodiment beingdescribed. In various embodiments, the fan 23 may be rotated at arelatively low speed at cruise to facilitate the low temperature rise.For example, the fan 23 may rotate at less than 2000 rpm, and/or mayhave a tip speed below Mn 1.1. The fan 23 of such embodiments may have afan diameter equal to or greater than 230 cm. The skilled person wouldappreciate that, in various embodiments, a gearbox 30 may be provided tofacilitate slower rotation of the fan 23, and that the engine cycle maybe designed around these parameters.

The aerodynamic design of the fan 23 may be selected to facilitateobtaining relatively low temperature rises across the fan as describedherein. For example, the fan 23 may be designed to have pressure ratiosat cruise of:

-   -   Fan Tip Pressure Ratio: in the range of 1.2-1.45; optionally in        the range of 1.35-1.44; and further optionally equal to 1.41;    -   Fan Root Pressure Ratio at cruise: in the range of 1.18-1.30,        and optionally equal to 1.24; and/or    -   Fan Pressure Ratio: in the range of 1.35-1.43, and optionally        equal to 1.39.

The fan tip pressure ratio is defined as the mean total pressure of theflow at the fan exit that subsequently flows (as flow B) through thebypass duct 22 to the mean total pressure at the inlet to the fan 23.With reference to FIGS. 4A and 4B, the mean total pressure of the flowat the fan exit that subsequently flows through the bypass duct 22 isthe mean total pressure over the surface that is immediately downstreamof the fan 23 and radially outside the streamsurface 110.

The fan root pressure ratio is defined as the mean total pressure of theflow at the fan exit that subsequently flows (as flow A) through theengine core 11 to the mean total pressure at the inlet to the fan 23.With reference to FIGS. 4A and 4B, the mean total pressure of the flowat the fan exit that subsequently flows through the engine core is themean total pressure of the flow that is immediately downstream of thefan 23 and radially inside the streamsurface 110. The mean totalpressure at the inlet to the fan 23 is the mean total pressure over thesurface that extends across the engine (for example from the hub 66 tothe tip 68 of the fan blade 66) and is immediately upstream of the fan23.

The fan pressure ratio is defined as the ratio of the mean totalpressure of the air flow at the exit of the fan 23 to the mean totalpressure of the air flow at the inlet of the fan 23.

The hub to tip ratio of the fan, defined as R_(Hub)/R_(fan tip), mayalso be selected as part of aerodynamic design considerations for thefan 23—for the engine 10 being described, the hub to tip ratio of thefan 23 is in the range from 0.285 to 0.2, and optionally in the rangefrom 0.24 to 0.27.

As mentioned above, the temperature across the trailing edge 64 b ofeach blade 64 generally varies with radius—the average temperature ofairflow B entering the bypass duct 22 is different from the averagetemperature of airflow A entering the engine core 11. A furthertemperature, T21, is therefore defined as the average temperature ofairflow A entering the engine core 11, which corresponds to the averagetemperature across the radially inner portion of the trailing edge 64 aof each fan blade 64 (the radially inner portion of each blade 64 beingthe remainder of the blade once the radially outer portion defined aboveis removed from consideration). T21 may be referred to as the inner fanrotor exit temperature or fan root exit temperature. T21 may be referredto as the core entry temperature, as it is the average temperature ofairflow entering the engine core 11 at cruise conditions. As thetemperature of the core airflow does not vary significantly between thetrailing edge 64 b of the fan blades 64 and the first stator/guide vane24 within the core 11, T21 may be measured anywhere within that regionas marked in FIG. 4B. For example, the core entry temperature (T21) maybe measured at/defined as any or all of the following:

(i) the temperature of the core airflow at the axial position of theforwardmost point 70 of the core casing 11 a (the splitter 70);

(ii) the temperature of the core airflow at the axial position of theleading edge of the forwardmost stator or rotor of the(forwardmost/lowest pressure) compressor 14; and/or

(iii) the temperature of the airflow across the trailing edge 64 b of aradially inner portion of each fan blade 64, the airflow across theradially inner portion of each fan blade 64 being arranged to providethe core airflow A.

The compressor exit temperature (T30) is defined as an averagetemperature of airflow at the exit from the compressor 15. T30 isdefined at the axial position of the trailing edge of the rearmost rotorof the compressor 15. In the embodiment being described, combustionequipment 16 located between the exit from the compressor 15 and theentrance to the turbine 17 located downstream of the compressor 15provides heat to the gas flow leaving the compressor 15, so increasingthe temperature of flow into the turbine 17 from T30 to T40—thedifference between T30 and T40 may be around 800 K, or more in someembodiments.

In the embodiment being described, the gas turbine engine comprises morethan one compressor 14, 15, and more specifically comprises twocompressors. In such embodiments, the compressor exit temperature T30 isdefined at the exit from the highest pressure compressor 15.

A core temperature rise may be defined as:

$\frac{{the}{compressor}{exit}{temperature}\left( {T30} \right)}{{the}{fan}{rotor}{entry}{temperature}\left( {T120{or}T20} \right)}.$

The core temperature rise may therefore measure the change in coreairflow A temperature caused by both the fan 23 and the compressor(s)14, 15. The core temperature rise is defined as the temperature ratioacross the core compression system; this may be thought of as a measureof core thermal efficiency. In the embodiments being described, the coretemperature rise is in the range from 3.1 to 4.0, and optionally in therange from 3.3 to 3.5. The core temperature rise may be equal to 3.33.

A core to fan tip temperature rise ratio may be defined as:

$\frac{{the}{core}{temperature}{rise}}{{the}{fan}{tip}{temperature}{rise}} = {\frac{T{30/T}120}{T{125/T}120} = \frac{T30}{T125}}$

As the temperature of both the core flow A and the bypass flow B is thesame at the leading edge 64 a of the fan blades 64 (T20=T120), the sametemperature value, T120, can be used for each flow A, B, so allowingT120 to cancel out as shown above.

The core to fan tip temperature rise ratio may be in the range from2.845 to 3.8, and optionally in the range from 2.9 to 3.2. The core tofan tip temperature rise ratio may be equal to 3 in some embodiments.The core to fan tip temperature rise ratio is therefore relatively highby virtue of the relatively high core temperature rise and relativelylow fan tip temperature rise. The engine cycle may be devised, and/orengine parameters selected, based on these parameters.

In the embodiment being described, a geared architecture and gearbox 30are used to facilitate the lower fan tip temperature rise as describedabove. In addition, a compressor design is selected to provide acompressor 14, 15 with an aerodynamic design that is efficient at a highlevel of loading. In the embodiments being described, the compressordesign comprises 13 or more stages of compression (including the fan 23as the first stage) so as to provide the desired effects. The compressordesign may comprise a maximum of 16 stages of compression (including thefan 23) in some such embodiments. Each stage may be defined as a rotor,or a rotor-stator pair. In the embodiment being described, with 13stages of compression—the fan 23 provides the first stage, the lowpressure compressor 14 provides the subsequent three stages, and thehigh pressure compressor 15 provides the final nine stages. Inalternative embodiments, the total number of compression stages mayvary, the number of compressors 14, 15 may vary, and/or the splitbetween the one or more compressor(s) may vary.

In the embodiment being described the compressors 14, 15 are axialcompressors. In various alternative embodiments, one or more of thecompressors 14, 15 may be a centrifugal compressor.

A high pressure ratio across the compressor(s) and an efficient level ofcore compression may therefore be achieved.

The core to fan tip temperature rise ratio may be thought of as arelationship between the temperature rise across the core compressionsystem (including the fan) and that across the bypass compression systemat the cruise operating conditions (i.e. the mid-cruise operating point,which is an altitude of 10700 m (35,000 ft), or optionally moreparticularly of 10668 m, and a speed of 0.85 Mn in the embodiment beingdescribed.

A core compressor temperature rise may be defined as:

$\frac{{the}{compressor}{exit}{temperature}\left( {T30} \right)}{{the}{core}{entry}{{temperature}\left( {T21} \right)}}$

The core compressor temperature rise may be in the range from 2.9 to4.0, and optionally in the range from 3.1 to 3.3. The core compressortemperature rise may be, for example, equal to 3.12.

A core compressor to fan tip temperature rise ratio may be defined as:

$\frac{{the}{core}{compressor}{temperature}{rise}}{{the}{fan}{tip}{temperature}{rise}} = {\frac{T{30/T}21}{T{125/T}120} = \frac{T120 \times T30}{T125 \times T21}}$

The core compressor to fan tip temperature rise ratio may be in therange from 2.67 to 3.8, and more specifically in the range from 2.67 to3.7 or from 2.67 to 3.5. The core compressor to fan tip temperature riseratio may be in the range from 2.80 to 2.95, and optionally may be equalto 2.81. The core compressor to fan temperature rise ratio may berelatively high, resulting from a relatively low fan tip temperaturerise, and/or a relatively high core compressor temperature rise.

As compared to the core temperature rise, the core compressortemperature rise excludes the temperature rise across the fan root 69,and therefore only measures the heat imparted to the airflow A by thecompressor(s) 14, 15.

In the embodiment being described, obtaining a relatively low fan tiptemperature rise is facilitated by use of a geared architecture and agearbox 30, allowing the fan 23 to rotate more slowly than otherdrivetrain components. The relatively high core compressor temperaturerise may be provided by having a core compressor aerodynamic design thatis efficient at a high level of loading, which typically can be achievedwith 13 stages of compression or greater as discussed above.

The core compressor to fan tip temperature rise ratio may be thought ofas a relationship between the temperature rise across the corecompression system (excluding the fan) and that across the bypasscompression system at the cruise operating condition.

The core compressor temperature rise is defined as the temperature ratioacross the core compression system; this may be thought of as a measureof core compressor thermal efficiency, by which the core compressorpressure rise is achieved.

The skilled person would appreciate that one or more of the followingengine features may be adjusted to obtain an engine 10 with a corecompressor to fan tip temperature rise ratio within the specified range:

-   -   A large flow area fan design, with the fan 23 being arranged to        rotate at a relatively slow speed (optionally facilitated by use        of a gearbox 30) in order to achieve a low fan tip temperature        rise; and/or    -   A core compression system with high levels of efficiency and        optimised loading that facilitate obtaining a high thermal        efficiency.        A low temperature rise across the fan root 69 may facilitate the        fan achieving a high propulsive efficiency while being operable        and mechanically feasible.

A fan root temperature rise may be defined as:

$\frac{{the}{core}{entry}{{temperature}\left( {T21} \right)}}{{the}{fan}{rotor}{entry}{temperature}\left( {T120} \right)}$

The fan root temperature rise may be in the range from 1.03 to 1.09, andoptionally in the range from 1.05 to 1.07. The fan root temperature risemay be, for example, equal to 1.07.

A core compressor to fan root temperature rise ratio may be definedusing the fan root temperature rise and the core compressor temperaturerise:

$\frac{{the}{core}{compressor}{temperature}{rise}}{{the}{fan}{root}{temperature}{rise}} = {\frac{T{30/T}21}{T{21/T}120} = \frac{T30 \times T120}{T21^{2}}}$

The core compressor to fan root temperature rise ratio may be in therange from 2.76 to 4.1, optionally in the range from 2.8 to 3.2, and maybe, for example, equal to 2.9.

The engine 10 may have a geared architecture comprising a gearbox 30.The engine 10 may have a high pressure ratio and an efficient level ofcore compression, for example achieved by having a core compressoraerodynamic design that is efficient at a high level of loading whichtypically can be achieved with 13 stages of compression or greater, asdiscussed above.

The fan root 69 may be designed to have a low temperature rise and a lowlevel of work to facilitate the operability of the fan and the obtainingof a high level of propulsive efficiency. The high level of propulsiveefficiency may be provided by a relatively straight fan root 69, havinga low level of curvature relative to the curvature of the fan tip. Forexample, the curvature of the fan root 69 may be less than 60% of thecurvature of the fan tip. In the embodiment being described, thecurvature of the root portion of the blade is between 40% and 60% lessthan the curvature across the tip portion of the blade, and optionallyaround 50% less. In alternative or additional embodiments, the curvatureof the root portion may be less than that of the tip portion by anamount within a range having a lower bound of any of 5%, 10%, 20%, 30%,40% and an upper bound of any of 40%, 50% or 60%. The listed percentagesare percentages of the blade camber (i.e. difference between a line thatis tangent to the camber line at the leading edge of the blade 68 and aline that is tangent to the camber line at the trailing edge of theblade 68). The skilled person would appreciate that the “root portion”of a fan blade is sometimes taken to mean the portion of a fan blade 64within the hub 66 and used to connect the blade 64 to the hub 66; thisis not the case as used herein—the root portion 69 refers to theradially inner portion of the blade as described elsewhere herein,extending from the hub 66 and across the entrance to the core 11. Theradially inner portion of the blade as defined herein may also be knownas the “hub section” of the blade by the skilled person.

The Overall Pressure Ratio (OPR) of an engine 10 with a core compressorto fan root temperature rise ratio in the listed range may be greaterthan 40.

In various embodiments, the engine core 11 comprises:

-   -   a first, lower pressure, turbine 19 (sometimes referred to as        the low pressure turbine, or LPT), a first compressor 14, and a        first core shaft 26 connecting the first turbine 19 to the first        compressor 14; and    -   a second, higher pressure, turbine 17 (sometimes referred to as        the high pressure turbine, or HPT), a second compressor 15, a        second core shaft 27 connecting the second turbine 17 to the        second compressor 15.

In such embodiments, a second turbine exit temperature (T42) may bedefined as an average temperature of airflow at the exit from the secondturbine 17 at cruise conditions and a first turbine exit temperature(T50) may be defined as an average temperature of airflow at the exitfrom the first turbine 19 at cruise conditions. T42 may be referred toas the high pressure turbine exit temperature. T50 may be referred to asthe low pressure turbine exit temperature.

In embodiments wherein the engine 10 comprises more than two turbines17, 19, and the highest pressure turbine 17 of the engine 10 may beselected as the second turbine 17 and the lowest pressure turbine 19 ofthe engine 10 may be selected as the first turbine 19.

As shown in FIG. 5A, T42 may be measured at the position of the rearmostrotor of the second turbine 17 and T50 may be measured at the positionof the rearmost rotor of the first turbine 19.

In the embodiment being described, the first turbine 19 is locatedimmediately downstream of the second turbine 17. The second turbine exittemperature T42 may therefore be measured anywhere between the secondturbine 17 and the first turbine 19, and may also be similar orequivalent to an entrance temperature (T44) for the first, lowerpressure, turbine 19. FIG. 5B illustrates the measurement region forT42. The skilled person would appreciate that there may be a change intemperature, for example of around 10-20 K, across this region, forexample due to cooling air. However, this difference may be too small asto noticeably affect the claimed ratios.

In embodiments with more than two turbines 17, 19, the highest pressureturbine exit temperature T42 may not be similar or equivalent to thelowest pressure turbine entry temperature T44.

The low(est) pressure turbine 19 is located immediately upstream of acore exit nozzle in the embodiment being described. T50 may be measuredat the position of the rearmost rotor of the first turbine 19, oranywhere within the core exit nozzle. The skilled person wouldappreciate that core gas stream temperature will gradually equilibratewith the surrounding atmosphere as it leaves/once it has left the engine10. FIG. 5B illustrates the measurement region for T50.

A low pressure turbine temperature change may be defined as:

$\frac{{the}{second}{turbine}{exit}{{temperature}{}\left( {T42} \right)}}{{the}{first}{turbine}{exit}{temperature}\left( {T50} \right)}.$

The low pressure turbine 19 temperature change may alternatively bedefined as follows, noting that the first turbine entrance temperatureT44 is generally comparable to the second turbine exit temperature T42:

$\frac{{the}{first}{turbine}{entrance}{temperature}\left( {T44} \right)}{{the}{first}{turbine}{exit}{temperature}\left( {T50} \right)}$

The skilled person would appreciate that temperature falls across aturbine in use, such that the entrance temperature (T42) to the first(low pressure) turbine 19 is higher than the exit temperature (T50) fromthe first turbine 19. The temperature change may therefore be describedas a temperature fall or temperature drop.

In embodiments with more than two turbines 17, 19, the low pressureturbine (LPT) temperature change may instead be called the lowestpressure turbine temperature change—providing a measure of thetemperature change over just the lowest pressure turbine.

The low pressure turbine temperature change therefore provides a measureof the average temperature change between the leading edge of the LPT 19nozzle guide vane 19 a (the stator of the forward-most rotor-stator pairof the LPT 19) and the trailing edge of the final rotor stage 19 c ofthe LPT 19, as indicated in FIG. 5A.

The low pressure turbine temperature change (T42/T50) may be in therange from 1.6 to 1.85, and optionally in the range from 1.65 to 1.8.The low pressure turbine temperature change may be, for example, equalto 1.68.

A turbine to fan tip temperature change ratio may be defined as shownbelow, using the fan tip temperature rise as defined above:

$\frac{{the}{low}{pressure}{turbine}{temperature}{change}}{{the}{fan}{tip}{temperature}{rise}}$

The turbine to fan tip temperature change ratio may be higher than inknown engines, for example being in the range from 1.46 to 2.0 andoptionally from 1.46 to 1.66. The turbine to fan tip temperature changeratio may be below 2.00, and optionally in the range from 1.5 to 1.8 insome embodiments.

In engines 10 of various embodiments with this temperature relationship,one or more of the following features may be present:

-   -   A gearbox that allows the LPT 19 to operate at a higher speed,        for example having a maximum operating speed between 4500 and        8000 rpm, e.g. for a fan 23 with a fan diameter in the range        from 330 cm to 380 cm or alternatively example having a maximum        operating speed between 7000 and 12000 rpm, e.g. for a fan 23        with a fan diameter in the range from 240 cm to 280 cm;        -   The skilled person would appreciate that, in embodiments            with a gearbox 30, LPT speed is generally equal to the fan            speed multiplied by the gear ratio of the gearbox 30.    -   A LPT 19 with favourable stage loading, for example having three        or more rotor stages;    -   An efficient aerodynamic fan design, for example having        relatively low fan root curvature as compared to fan tip        curvature (as described in more detail elsewhere herein);    -   A fan 23 arranged to rotate at a relatively slow speed at        cruise, optionally enabled by a gearbox 30;        -   The rotational speed of the fan at cruise conditions may be,            for example, less than 2500 rpm, or less than 2300 rpm.        -   For an engine having a fan diameter in the range of from 220            cm to 300 cm (for example 240 cm to 280 cm or 250 cm to 270            cm), the fan speed may be in the range of from 1700 rpm to            2500 rpm, for example in the range of from 1800 rpm to 2300            rpm, for example in the range of from 1900 rpm to 2100 rpm.        -   For an engine having a fan diameter in the range of from 330            cm to 380 cm, the fan speed may be in the range of from 1200            rpm to 2000 rpm, for example in the range of from 1300 rpm            to 1800 rpm, for example in the range of from 1400 rpm to            1800 rpm.            and/or    -   An engine designed to be low specific thrust, high bypass ratio        and to have a low fan pressure ratio (the ratio of the mean        total pressure of the air flow at the exit of the fan 23 to the        mean total pressure of the air flow at the inlet of the fan 23,        at cruise conditions).        -   “Low specific thrust” may mean, in various embodiments, a            thrust at cruise conditions in the range of 60 to 100            NKg⁻¹s, and optionally in the range of 70-90 NKg⁻¹s. For            example, the thrust at cruise may be equal to or below 90            NKg⁻¹s, and optionally equal to or below 88 or 85 NKg⁻¹s.        -   “High bypass ratio” may mean, in various embodiments, a            bypass ratio at cruise conditions in the range of, for            example, 12.5 to 30;        -   “Low fan pressure ratio” may mean, in various embodiments, a            fan pressure ratio at cruise conditions in the range of 1.2            to 1.45; and optionally in the range 1.35 to 1.43. Further            optionally, the fan pressure ratio at cruise may be in the            range 1.37 to 1.41 or 1.38 to 1.40. For example, the fan            pressure ratio at cruise may be equal to 1.39. In            alternative or additional embodiments, the fan pressure            ratio at cruise may be equal to or around 1.35, 1.36, 1.37,            1.38, 1.39, 1.40, 1.41, 1.42 or 1.43, and may for example be            in the range from 1.39 to 1.43, from 1.35 to 1.40, and/or            from 1.37 to 1.40.

To facilitate the operation of the lower pressure turbine (LPT) 19 so asto provide a larger temperature change (a higher magnitude drop intemperature across the lower pressure turbine 19), a gearbox 30 may beprovided to allow the LPT 19 to operate at a favourable, higher, speed.

The LPT 19 may be designed to provide a favourable stage loading for thelarger temperature change. Although dependent on engine thrust, such adesign may typically be achieved with a LPT with three or more rotorstages, and optionally with four or more rotor stages.

In an engine core 11 comprising a second, higher pressure, turbine 17, asecond compressor 15, a second, core shaft 27 connecting the secondturbine 17 to the second compressor 15, and a first, lower pressure,turbine 19, a first compressor 14, and a first core shaft 26 connectingthe first turbine 19 to the first compressor 14, a second turbineentrance temperature (T40) may be defined as an average temperature ofairflow at the entrance to the second turbine 17 at cruise conditions.T40 may be measured/defined at the upstream (leading) edge of an inletnozzle guide vane 17 a of the high pressure turbine 17, as shown in FIG.5A. An inlet nozzle guide vane 17 a may be thought of as theforward-most stator of the high pressure turbine 17. In the embodimentbeing described, the gas stream temperature increases between the outletof the compressor 15 and the inlet of the turbine 17 due to thecombustor equipment 16; T40 may therefore be measured, or determined fora point, anywhere between the exit from the combustor equipment 16 andthe entrance to the second turbine 17, as marked in FIG. 5B, and maymore specifically be measured or otherwise determined at a leading edgeof the most upstream stator 17 a of the second turbine 17.

A high(est) pressure turbine temperature change may be defined as:

$\frac{{the}{second}{turbine}{entrance}{temperature}\left( {T40} \right)}{{the}{second}{turbine}{exit}{temperature}\left( {T42} \right)}.$

The high pressure turbine (HPT) temperature change may therefore bedefined as the average temperature change between upstream of the HPT 17inlet nozzle guide vane 17 a (the forward most stator of the HPT 17) andafter the final rotor stage of the HPT 17 as indicated in FIG. 5A.

In embodiments with more than two turbines 17, 19, the high pressureturbine temperature change may instead be referred to as the highpressure turbines temperature change and may be a measure of thetemperature change across all turbines excluding the lowest pressureturbine 19. If the second turbine 17 is the highest pressure turbine,the second turbine exit temperature T42 may therefore be replaced withthe lowest-but-one pressure turbine exit temperature in the calculationof the ratio, which may be at least substantially equal to the lowestpressure turbine entry temperature T44. T44 may not be similar to orequal to T42 in such embodiments. The high pressure turbine temperaturechange may be in the range from 1.40 to 1.55, and optionally in therange from 1.44 to 1.52. The high pressure turbine temperature changemay be equal to 1.5, for example being 1.50 or 1.51.

A low to high pressure turbine temperature change ratio may then bedefined as below, noting that T42 is generally at least similar to T44(e.g. within 10-20 K):

$\frac{{the}{low}{pressure}{turbine}{temperature}{change}}{{the}{high}{pressure}{turbine}{temperature}{change}} = {\frac{T{44/T}50}{T{40/T}42} = {\frac{T42 \times T44}{T50 \times T40} \approx \frac{T42^{2}}{T50 \times T40}}}$

The low to high pressure turbine temperature change ratio, which mayalso be referred to as a temperature fall ratio, may be in the rangefrom 1.09 to 1.30, and optionally in the range from 1.10 to 1.25.

The low to high pressure turbine temperature change ratio provides arelationship between the temperature change across the low pressureturbine 19 and the temperature change across the high pressure turbine17 at the cruise operating conditions.

To reduce fuel burn, and optionally reduce or minimise core size, and/ormaximise thermal efficiency across the high pressure turbine, theinventors appreciated that a relatively low temperature change acrossthe higher pressure turbine (HPT) 17 (as compared to the temperaturechange across the lower pressure turbine 19) may be beneficial.

In various embodiments, this relatively low HPT temperature change maybe obtained by using an HPT 17 with an efficient design, for examplehaving two rotor stages, or only a single rotor stage.

In the embodiments being described, the first turbine 19 is arranged toreceive airflow from the exit of the second turbine 17, such that thefirst turbine entrance temperature (T42) is generally similar to thesecond turbine exit temperature (T42), sometimes with a 10 to 20 Kdifference due to the introduction of cooling air.

In the embodiments being described, the second turbine 17 is arranged toreceive airflow from the exit of the (high pressure) compressor 15; thisairflow passes via the combustion equipment 16 between the compressorand the turbine, such that T30 (the second turbine entrance temperature)is higher than the compressor exit temperature T40.

In engines 10 of various embodiments with this temperature relationship,one or more of the following features may be present:

-   -   A gearbox arranged to allow the LPT 19 to operate at a        favourable, higher, speed;    -   A LPT 19 with optimal stage loading, for example having three or        more rotor stages 19 a, 19 b, 19 c;    -   A high pressure compressor 15 with an aerodynamic design and a        low level of loading, for example having nine or more rotor        stages;    -   An efficient HPT 17, for example having two rotor stages or        fewer.

The present disclosure also relates to methods 1000 of operating a gasturbine engine 10 on an aircraft 50. The methods 1000 are illustrated inFIG. 8 . The method 1000 comprises starting up 1002 the engine 10 (e.g.prior to taxiing on a runway), and operating 1004 the engine duringtaxiing, take-off, and climb of the aircraft 50, as suitable, so as toreach cruise conditions. Once cruise conditions have been reached, themethod 1000 then comprises operating 1006 the gas turbine engine 10described in embodiments elsewhere herein to provide propulsion undercruise conditions.

The gas turbine engine 10 is operated such that any one or more of theparameters or ratios defined herein are within the specified ranges. Forexample, the method comprises operating 1006 the gas turbine engine 10such that any one or more of:

a) the fan hub to tip ratio of:

$\frac{{the}{fan}{hub}{radius}(103)}{{the}{fan}{tip}{radius}(102)}$

is in the range from 0.2 to 0.285; and

the fan tip temperature rise of:

$\frac{{t{he}}{fan}{tip}{rotor}{exit}{temperature}\left( {T125} \right){in}{Kelvin}}{{the}{fan}{rotor}{entry}{temperature}\left( {T120} \right){in}{Kelvin}}$

is in the range from 1.11 to 1.05;

b) the core to fan temperature rise ratio of:

$\frac{{the}{core}{temperature}{rise}}{{the}{fan}{tip}{temperature}{rise}}$

is in the range from 2.845 to 3.8;

c) the core compressor to fan tip temperature rise ratio of:

$\frac{{the}{core}{compressor}{temperature}{rise}}{{the}{fan}{tip}{temperature}{rise}}$

is in the range from 2.67 to 3.8, and optionally 2.67 to 3.7;

d) the core compressor to fan root temperature rise ratio of:

$\frac{{the}{core}{compressor}{temperature}{rise}}{{the}{fan}{root}{temperature}{rise}}$

is in the range from 2.76 to 4.1.

e) the turbine to fan tip temperature change ratio of:

$\frac{{the}{low}{pressure}{turbine}{temperature}{change}}{{the}{fan}{tip}{temperature}{rise}}$

is in the range from 1.46 to 2.0; and/or

f) a low to high pressure turbine temperature change ratio of:

$\frac{{the}{low}{pressure}{turbine}{temperature}{change}}{{the}{high}{pressure}{turbine}{temperature}{change}}$

is in the range from 1.09 to 1.30, and optionally from 1.10 to 1.25.

FIG. 10 illustrates an example aircraft 50 having a gas turbine engine10 attached to each wing 52 a, 52 b thereof. Each gas turbine engine 10is attached via a respective pylon 54 a, 54 b. When the aircraft 50 isflying under cruise conditions, as defined herein, each gas turbineengine 10 operates according to the parameters defined herein. Forexample, the gas turbine engines 10 operate such that any one or more ofthe conditions (a) to (f) defined for the method 1000 above areobtained.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

1. A gas turbine engine for an aircraft comprising: an engine corecomprising a turbine, a compressor, and a core shaft connecting theturbine to the compressor, wherein a compressor exit temperature isdefined as an average temperature of airflow at the exit from thecompressor at cruise conditions and a core entry temperature is definedas an average temperature of airflow entering the engine core at cruiseconditions, and a core compressor temperature rise is defined as:$\frac{{the}{compressor}{exit}{temperature}}{{the}{core}{entry}{temperature}};$the engine core further comprising an annular splitter at which the flowis divided between a core flow that flows through the engine core, and abypass flow that flows along a bypass duct, wherein stagnationstreamlines around the circumference of the engine, stagnating on aleading edge of the annular splitter, form a streamsurface forming aradially inner boundary of a streamtube that contains all of the bypassflow; and a fan located upstream of the engine core, the fan comprisinga plurality of fan blades extending from a hub, each fan blade having aleading edge and a trailing edge, each fan blade having a radially outerportion lying within the streamtube that contains the bypass flow, andwherein a fan rotor entry temperature is defined as an averagetemperature of airflow across the leading edge of each fan blade atcruise conditions and a fan tip rotor exit temperature is defined as anaverage temperature of airflow across the radially outer portion of eachfan blade at the trailing edge at cruise conditions and a fan tiptemperature rise is defined as:$\frac{{the}{fan}{tip}{rotor}{exit}{temperature}}{{the}{fan}{rotor}{entry}{temperature}},$and wherein the gas turbine engine is configured to have a corecompressor to fan tip temperature rise ratio of:$\frac{{the}{core}{compressor}{temperature}{rise}}{{the}{fan}{tip}{temperature}{rise}}$in the range from 2.67 to 3.8, the compressor exit temperature in therange of from 750 K to 1050 K, and wherein the cruise conditions meansthe conditions at mid-cruise of the aircraft to which the gas turbineengine is attached, and the cruise conditions at mid-cruise means theconditions experienced by the aircraft and the gas turbine engine at apoint between top of climb and start of descent at which 50% of thetotal fuel that is burned between the top of climb and the start ofdescent has been burned.
 2. The gas turbine engine according to claim 1,wherein the core compressor to fan tip temperature rise ratio is in therange from 2.67 to 3.7.
 3. The gas turbine engine according to claim 1,wherein the core compressor to fan tip temperature rise ratio is in therange from 2.80 to 2.95.
 4. The gas turbine engine according to claim 1,wherein the compressor exit temperature in the range of from 780 K to1000 K.
 5. The gas turbine engine according to claim 1, wherein the corecompressor temperature rise is in the range from 2.9 to 4.0.
 6. The gasturbine engine according to claim 1, wherein the core compressortemperature rise is in the range from 3.1 to 3.3.
 7. The gas turbineengine according to claim 1, wherein the engine comprises more than onecompressor, and wherein the compressor exit temperature is measured atthe exit from the highest pressure compressor.
 8. The gas turbine engineaccording claim 1, wherein a fan tip radius of the fan is definedbetween a centreline of the engine and an outermost tip of each fanblade at its leading edge and a fan hub radius of the fan is definedbetween the centreline of the engine and an outer surface of the hub atthe radial position of the leading edge of each fan blade, wherein a fanhub to tip ratio of:$\frac{{the}{fan}{hub}{radius}}{{the}{fan}{tip}{radius}}$ is in therange from 0.2 to 0.285.
 9. The gas turbine engine according to claim 1,wherein a specific thrust of the engine at cruise conditions, defined asnet engine thrust divided by mass flow rate through the engine, is inthe range from 50 to 100 Nkg⁻¹s.
 10. The gas turbine engine according toclaim 1, wherein the fan tip temperature rise is in the range from 1.05to 1.11.
 11. The gas turbine engine according to claim 1, wherein a fanroot temperature rise defined as:$\frac{{the}{core}{entry}{temperature}}{{the}{fan}{rotor}{entry}{temperature}}$is in the range from 1.03 to 1.09.
 12. The gas turbine engine accordingto claim 11, wherein the fan root temperature rise is in the range from1.05 to 1.07.
 13. The gas turbine engine according to claim 11, whereina core compressor to fan root temperature rise ratio of$\frac{{the}{core}{compressor}{temperature}{rise}}{{the}{fan}{root}{temperature}{rise}}$is in the range from 2.76 to 4.10.
 14. The gas turbine engine accordingto claim 13, wherein the core compressor to fan root temperature riseratio is in the range from 2.8 to 3.2.
 15. The gas turbine engineaccording to claim 1, wherein: the turbine is a second turbine, thecompressor is a second compressor, and the core shaft is a second coreshaft; the engine core further comprises a first turbine, a firstcompressor, and a first core shaft connecting the first turbine to thefirst compressor; and the second turbine, second compressor, and secondcore shaft are arranged to rotate at a higher rotational speed than thefirst core shaft.
 16. The gas turbine engine according to claim 15,wherein a first turbine entrance temperature is defined as an averagetemperature of airflow at the entrance to the first turbine at cruiseconditions and a first turbine exit temperature is defined as an averagetemperature of airflow at the exit from the first turbine at cruiseconditions, and wherein a first turbine temperature change defined as:$\frac{{the}{first}{turbine}{entrance}{temperature}}{{the}{first}{turbine}{exit}{temperature}}$is in the range from 1.6 to 1.85.
 17. The gas turbine engine accordingto claim 16, wherein a turbine to fan tip temperature change ratiodefined as:$\frac{{the}{first}{turbine}{temperature}{change}}{{the}{fan}{tip}{temperature}{rise}}$is in the range from 1.46 to 2.0.
 18. The gas turbine engine accordingto claim 1, wherein: (i) the fan tip radius is in the range from 110 cmto 150 cm; or (ii) the fan tip radius is in the range from 155 cm to 200cm.
 19. The gas turbine engine according to claim 1, further comprisinga gearbox that receives an input from the core shaft and outputs driveto the fan so as to drive the fan at a lower rotational speed than thecore shaft, and wherein the gearbox has a gear ratio in the range offrom 3.2 to
 5. 20. A method of operating a gas turbine engine on anaircraft, wherein the method comprises: operating the gas turbine engineto provide propulsion under cruise conditions according to claim 1.